Midshaft rating for turbomachine engines

ABSTRACT

A turbomachine engine includes a fan section having a fan shaft, and a core engine having one or more compressor sections, one or more turbine sections that includes a power turbine, and a combustion chamber in flow communication with the compressor sections and turbine sections. The turbomachine engine includes a low-speed shaft coupled to the power turbine and having a midshaft that extends from a forward bearing to an aft bearing. The low-speed shaft is characterized by a midshaft rating (MSR) between two hundred (ft/sec) 1/2  and three hundred (ft/sec) 1/2 . The low-speed shaft has a redline speed between fifty and two hundred fifty feet per second (ft/sec). The turbomachine engine includes a gearbox assembly that couples the fan shaft to the low-speed shaft and characterized by a gearbox assembly mode less than 95% of a midshaft mode of the midshaft or greater than 105% of the midshaft mode.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a continuation-in-part application of U.S. patentapplication Ser. No. 17/328,795, filed May 24, 2021, the entire contentsof which is incorporated by reference in its entirety.

This application is related to U.S. patent application Ser. No.17/328,800, filed May 24, 2021, U.S. patent application Ser. No.18/058,040, filed Nov. 22, 2022, and U.S. patent application Ser. No.18/058,036, filed Nov. 22, 2022. The entire contents of theaforementioned applications are incorporated by reference in theirentireties.

TECHNICAL FIELD

This application is generally directed to turbomachine engines,including turbomachine shafts, and a method of driving such turbomachineshafts in such turbomachine engines.

BACKGROUND

A turbofan engine, or turbomachinery engine, includes one or morecompressors, and a power turbine that drives a bypass fan. The bypassfan is coupled to the power turbine via a turbomachine shaft.

BRIEF DESCRIPTION OF THE DRAWINGS

The foregoing and other features and advantages will be apparent fromthe following, more particular, description of various embodiments, asillustrated in the accompanying drawings, wherein like reference numbersgenerally indicate identical, functionally similar, and/or structurallysimilar elements.

FIG. 1 shows a schematic, cross-sectional view of a ducted, direct-drivegas turbine engine.

FIG. 2 shows a schematic, cross-sectional view of a ducted,indirect-drive gas turbine engine.

FIG. 3 shows a schematic view of an unducted, three-stream gas turbineengine.

FIG. 4 shows an enlarged view of a portion of the cross-sectional viewof FIG. 1 .

FIG. 5A shows a cross-sectional view of a steel shaft.

FIG. 5B shows a cross-sectional view of a composite shaft.

FIG. 6A shows a cross-sectional view of a uniform shaft with a constantdiameter and thickness.

FIG. 6B shows a cross-sectional view of a concave shaft with a constantdiameter and a variable thickness.

FIG. 6C shows a cross-sectional view of a convex shaft with a variablediameter and a variable thickness.

FIG. 7A shows a schematic view of a shaft using a four-bearing straddleconfiguration.

FIG. 7B shows a schematic view of a shaft using a four-bearing outboundconfiguration.

FIG. 7C shows a schematic view of a shaft using an inbound duplexconfiguration.

FIG. 7D shows a schematic view of a shaft using an outbound duplexconfiguration.

FIG. 7E shows a schematic view of a shaft using a two-bearingconfiguration.

FIG. 8A shows an enlarged, schematic, cross-sectional view of a gearboxassembly of a gas turbine engine, taken along a centerline axis of thegas turbine engine.

FIG. 8B shows a schematic, cross-sectional view of the gearbox assemblyof FIG. 8A, translated into a representative vibratory system.

FIG. 9A shows an enlarged, schematic, cross-sectional view of a gearboxassembly of a gas turbine engine, taken along a centerline axis of thegas turbine engine, according to another embodiment.

FIG. 9B shows a schematic, cross-sectional view of the gearbox assemblyof FIG. 9A, translated into a representative vibratory system.

FIG. 10 shows a schematic, cross-sectional view of a gearbox assembly ofa gas turbine engine with an oil transfer device.

FIG. 11A shows a schematic view of the degrees of freedom of lateralstiffness.

FIG. 11B shows a schematic view of the degrees of freedom of bendingstiffness.

FIG. 11C shows a schematic view of the degrees of freedom of torsionalstiffness.

FIG. 12A shows an enlarged, schematic, cross-sectional side view of agearbox assembly with a mounting assembly for a gas turbine engine,taken at a centerline axis of the gas turbine engine.

FIG. 12B shows an enlarged, schematic, partial cross-sectional side viewof a portion of the mounting assembly of FIG. 12A.

FIG. 12C shows a schematic, partial cross-sectional view of a ring gearassembly for the gearbox assembly of FIG. 12A.

FIG. 12D shows an enlarged, schematic, partial cross-sectional view ofthe mounting assembly of FIG. 12C.

FIG. 12E shows an enlarged, schematic side view of the mounting assemblyof FIG. 12C.

FIG. 13 shows an enlarged, schematic cross-sectional side view of agearbox assembly with a mounting assembly for a gas turbine engine,taken at a centerline axis of the gas turbine engine, according toanother embodiment.

FIG. 14 shows an enlarged, schematic partial cross-sectional side viewof a portion of a mounting assembly, according to another embodiment.

FIG. 15 shows an enlarged, schematic partial cross-sectional view of aring gear assembly for a gas turbine engine.

FIG. 16A shows a schematic, cross-sectional side view of a gearboxassembly for a gas turbine engine, taken along a centerline axis of thegas turbine engine.

FIG. 16B shows a schematic view of a planet gear with a first stageplanet gear and a second stage planet gear.

FIG. 17 shows a schematic view of a planet gear having a single stagefor a gearbox assembly.

FIG. 18 shows an enlarged, schematic side cross-sectional view of agearbox assembly with a mounting assembly for a gas turbine engine,taken at a centerline axis of the gas turbine engine.

FIG. 19A shows a plot depicting a range of a midshaft rating relative toa range of outer diameter redline speeds.

FIG. 19B shows a plot depicting a range of a midshaft rating relative toa range of length-diameter ratios.

DETAILED DESCRIPTION

Additional features, advantages, and embodiments of the presentdisclosure are set forth or apparent from a consideration of thefollowing detailed description, drawings, and claims. Moreover, both theforegoing summary of the present disclosure and the following detaileddescription are exemplary and intended to provide further explanationwithout limiting the scope of the disclosure as claimed.

Various embodiments are discussed in detail below. While specificembodiments are discussed, this is done for illustration purposes only.A person skilled in the relevant art will recognize that othercomponents and configurations may be used without departing from thespirit and the scope of the present disclosure.

As used herein, the terms “first,” “second,” “third,” and “fourth” maybe used interchangeably to distinguish one component from another andare not intended to signify location or importance of the individualcomponents.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine or vehicle, and refer to the normal operational attitudeof the gas turbine engine or vehicle. For example, with regard to a gasturbine engine, forward refers to a position closer to an engine inletand aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The terms “coupled,” “fixed,” “attached to,” and the like, refer to bothdirect coupling, fixing, or attaching, as well as indirect coupling,fixing, or attaching through one or more intermediate components orfeatures, unless otherwise specified herein.

The term “propulsive system” refers generally to a thrust-producingsystem, which thrust is produced by a propulsor, and the propulsorprovides the thrust using an electrically-powered motor(s), a heatengine such as a turbomachine, or a combination of electrical motor(s)and a turbomachine.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

As used herein, the terms “axial” and “axially” refer to directions andorientations that extend substantially parallel to a centerline of theturbine engine. Moreover, the terms “radial” and “radially” refer todirections and orientations that extend substantially perpendicular tothe centerline of the turbine engine. In addition, as used herein, theterms “circumferential” and “circumferentially” refer to directions andorientations that extend arcuately about the centerline of the turbineengine.

As used herein, “redline speed” means the maximum expected rotationalspeed of a shaft during normal operation of an engine. The redline speedmay be expressed in terms of rotations per second in Hertz (Hz),rotations per minute (RPM), or as a linear velocity of the outerdiameter of the shaft in terms of feet per second. For a gas turbineengine that has a high-speed shaft and a low-speed shaft, both thehigh-speed shaft and the low-speed shaft have redline speeds.

As used herein, “critical speed” means a rotational speed of the shaftthat is about the same as the fundamental, or natural frequency of afirst-order bending mode of the shaft (e.g., the shaft rotates at eightyHz and the first-order modal frequency is eighty Hertz). When the shaftrotates at the critical speed, the shaft is expected to have a maximumamount of deflection, hence instability, due to excitation of thefirst-order bending mode of the shaft. The critical speed may beexpressed in terms of rotations per second in Hertz (Hz), rotations perminute (RPM), or as a linear velocity of the outer diameter of the shaftin terms of feet per second.

As used herein, “critical frequency” and “fundamental frequency” arereferred to interchangeably and refer to the fundamental, or naturalfrequency, of the first-order bending mode of the shaft.

The term “subcritical speed” refers to a shaft redline speed that isless than the fundamental, or natural frequency of the first-orderbending mode of the shaft (e.g., the shaft rotates at a redline speed of70 Hz while the first-order modal frequency is about 80 Hertz). When therotational speed is subcritical the shaft is more stable than whenrotating at a critical speed. A “subcritical shaft” is a shaft that hasa redline speed below the critical speed of the shaft.

The term “supercritical speed” refers to a shaft rotational speed thatis above the fundamental, or natural frequency of the first-orderbending mode of the shaft (e.g., the shaft rotates at eighty Hz whilethe first-order modal frequency is about seventy Hertz). A supercriticalshaft is less stable than a subcritical shaft because the shaft speedcan pass through the critical speed since its fundamental mode is belowthe redline speed. A “supercritical shaft” is a shaft that has a redlinespeed above the critical speed of the shaft.

The term “critical modal frequency” for a gearbox or FGBx is a naturalfrequency of vibration for a gearbox assembly, characterized by modalproperties (mode shape, strain energy in various supporting structural,etc.) producing lateral inertial bending or displacement reaction forcesthrough a gearbox sun-gear—midshaft coupling when the gearbox assembly.The gearbox assembly can produce a significant dynamic responsecharacterized by these modal properties when there is an external forceapplied with an excitation frequency at or near FGBX through the sungear—midshaft coupling (e.g. by motion of the midshaft) or when thegearbox assembly undergoes a periodic acceleration at or near FGBx.

The terms “low” and “high,” or their respective comparative degrees(e.g., “lower” and “higher”, where applicable), when used with thecompressor, turbine, shaft, or spool components, each refers to relativepressures and/or relative speeds within an engine unless otherwisespecified. For example, a “low-speed shaft” defines a componentconfigured to operate at a rotational speed, such as a maximum allowablerotational speed, which is lower than that of a “high-speed shaft” ofthe engine. Alternatively, unless otherwise specified, theaforementioned terms may be understood in their superlative degree. Forexample, a “low-pressure turbine” may refer to the lowest maximumpressure within a turbine section, and a “high-pressure turbine” mayrefer to the highest maximum pressure within the turbine section. Theterms “low” or “high” in such aforementioned regards may additionally,or alternatively, be understood as relative to minimum allowable speedsand/or pressures, or minimum or maximum allowable speeds and/orpressures relative to normal, desired, steady state, etc., operation ofthe engine.

The terms “lateral stiffness” and “lateral structural stiffness” areused interchangeably and refer to the stiffness (force per unit length)of a component when the component is displaced, e.g., one inch, in therespective lateral direction. That is, the stiffness of a component inthe radial direction (direction Y in FIGS. 1 to 3 and 11A) and thelateral direction (direction X in FIG. 11A; into and out of the page inFIGS. 1 to 3 ) would be, e.g., pounds per inch displacement in the Y andX directions. The lateral stiffness as defined as shown in FIG. 11A. Thelateral stiffness is identified herein as K^(L).

The terms “bending stiffness” and “bending structural stiffness” areused interchangeably and refer to the stiffness (force-length/radian) ofa component when displaced by one radian in the pitch or yaw planes.That is, the stiffness of a component in the pitch direction (rotationwithin the Y and Z plane in FIGS. 1 to 3 and 11B) and the yaw direction(rotation within the Z and X plane in FIGS. 1 to 3 and 11B) would bee.g., pounds-inch per radian rotation. The bending stiffness as definedas shown in FIG. 11B. The bending stiffness is identified herein asK^(B).

The term “casing” herein refers to the structure that defines an airflowpath (e.g., wall of duct, or casing). A mounting to the casing may be adirect bolted connection or through a load bearing frame.

A “static structure” as herein referred means any structural part of anengine that is non-rotating.

The terms “torsional stiffness” and “torsional structural stiffness” areused interchangeably and refer to the stiffness (force-length/radian) ofa component when displaced by one radian in the X-Y plane (within the Xand Y plane in FIG. 11C, or about a plane parallel to the LP midshaftcenterline). The torsional stiffness is defined as shown in FIG. 11C.The torsional stiffness herein is identified as K^(T).

The term “lateral damping” refers to the structural damping of acomponent in the lateral direction at a frequency of vibration. Thelateral damping is identified herein as C^(L).

The term “bending damping” refers to the structural damping of acomponent in the bending direction at a frequency of vibration. Thebending damping is identified herein as C^(B).

The term “torsional damping” refers to the structural damping of acomponent in the torsional or rotational direction at a frequency ofvibration. The torsional damping is identified herein as C^(T).

As used herein, a “flex coupling” is a mounting structure, such as ashaft, that connects a gearbox assembly to a low-speed shaft of the gasturbine engine. A flex coupling allows displacement of the gearboxassembly with respect to the low-speed shaft in the axial direction, theradial direction, and/or the circumferential direction, and a stiffnessand damping of the flex coupling can be tuned to achieve a desiredvibrational response through the flex coupling. An example of a flexcoupling 845 is shown in FIG. 8A and detailed further below.

As used herein, a “flex mount” is a mounting structure couples thegearbox assembly to an engine static structure of the turbomachineengine. A flex mount allows displacement of the gearbox assembly withrespect to the engine static structure in the axial direction, theradial direction, and/or the circumferential direction, and a stiffnessand damping of the flex coupling can be tuned to achieve a desiredvibrational response through the flex coupling. An example of a flexmount 847 is shown in FIG. 8A and detailed further below.

As used herein, a “deflection limiter” is a component that is used tomechanically limit the maximum deflection of the gearbox assembly beyonda threshold level. A deflection limiter can limit deflections in theradial direction, in the axial direction, and/or in the circumferentialdirection. Examples of deflection limiters are provided in FIGS.12B,12E, 14 , and 15, as detailed further below.

As used herein, a “damper” is a device or a component that absorbsvibrations emanating from the gearbox assembly and/or from the midshaft.A damper as used herein can include devices or components for lateraldamping, bending damping, and/or torsional damping. Examples of dampersare provided in FIGS. 8B, 9B, 12B, 12D, 12E, 14, and 15 .

As used herein, a “helical gear” is a cylindrical gear with gear teeth,also referred to as helical gear teeth, that extend at an angle (a helixangle) to an axis of rotation (e.g., in the axial direction) of thegear.

As used herein, a “bihelical gear” is a gear with two sets of helicalgear teeth that oppositely disposed with respect to each other at ahelix angle with respect to the axis of rotation.

As used herein, a “helix axis” is an axis along a length of a geartooth. For example, the helix axis 1767 (FIG. 17 ) is normal to an endface of the gear tooth. The helix axis 1767 of a helical gear isdisposed at a helix angle R with respect to the axis 1713 (FIG. 17 ) ofthe gear (e.g., the axis of rotation).

Here and throughout the specification and claims, range limitations arecombined, and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

One or more components of the turbomachine engine described herein belowmay be manufactured or formed using any suitable process, such as anadditive manufacturing process, such as a three-dimensional (3D)printing process. The use of such a process may allow such a componentto be formed integrally, as a single monolithic component, or as anysuitable number of sub-components. In particular, the additivemanufacturing process may allow such a component to be integrally formedand include a variety of features not possible when using priormanufacturing methods. For example, the additive manufacturing methodsdescribed herein enable the manufacture of shafts having uniquefeatures, configurations, thicknesses, materials, densities,passageways, headers, and mounting structures that may not have beenpossible or practical using prior manufacturing methods. Some of thesefeatures are described herein.

This disclosure and various embodiments relate to a turbomachineryengine, also referred to as a turbine engine, a gas turbine engine, aturboprop engine, or a turbomachine. These turbomachinery engines can beapplied across various technologies and industries. Various embodimentsmay be described herein in the context of aeronautical engines andaircraft machinery.

In some instances, a turbomachinery engine is configured as a directdrive engine. In other instances, a turbomachinery engine can beconfigured as an indirect drive engine with a gearbox. In someinstances, a propulsor of a turbomachinery engine can be a fan encasedwithin a fan case and/or nacelle. This type of turbomachinery engine canbe referred to as “a ducted engine.” In other instances, a propulsor ofa turbomachinery engine can be exposed (e.g., not within a fan case or anacelle). This type of turbomachinery engine can be referred to as “anopen rotor engine” or an “unducted engine.”

A turbofan engine, or turbomachinery engine, includes a core engine anda power turbine that drives a bypass fan. The bypass fan is coupled tothe core engine via a turbomachine shaft. The bypass fan generates themajority of the thrust of the turbofan engine. The generated thrust canbe used to move a payload (e.g., an aircraft). A turbomachine shaftcoupled to the power turbine and fan (either directly or through agearbox) can experience vibrations during operation of the engine (e.g.,during rotation of the shaft). For example, when the shaft rotates atits critical speed, the shaft will vibrate excessively. The excessivevibration is due primarily to excitation of a first-order beam bendingmode of the shaft. Thus, the shaft may be characterized by itsfirst-order bending mode, the fundamental resonance frequency(fundamental frequency) of this mode, and the shaft's critical speed ofrotation. If the first-order bending mode may be excited by a low-speedshaft rate occurring during a standard operating range of the engine,undetected vibration as well as an increased risk of whirl instability,may result. There is a continuing need to address vibrations induced byrotating shafts in turbomachinery engines.

Newer engine architectures may be characterized by faster shaft speedsfor the low-pressure turbine (LPT), and longer shafts to accommodate alonger engine core (e.g., the high-pressure compressor, the combustor,and the high-pressure turbine). Additionally, it is desirable to housethe engine core within a smaller space. These trends can result inreductions in stiffness-to-weight ratio for the shaft and structure thatinfluence dynamics of the LP shaft, which may have the effect oflowering the critical speed and/or limiting the available options forincreasing the critical speed for the LPT's shaft (referred to as thelow-speed shaft or the low-pressure (LP) shaft). Accordingly, differentapproaches for engine types, midshaft geometry, bearing support, andmaterial compositions are required for next-generation turbomachineengines, to permit high-speed operation without resulting in an unstablebending mode during regular operation. The inventors, tasked withfinding a suitable design to meet these requirements, conceived andtested a wide variety of shafts having different combinations ofstiffness, material, bearing type and location, shaft length, anddiameter in order to determine which embodiment(s) were most promisingfor a variety of contemplated engine designs. The various embodiments,as described herein and as shown in the figures, include turbomachineshafts that employ one or more of these techniques to increase thecritical speed of the first-order bending mode.

FIG. 1 shows a schematic, cross-sectional view of a ducted,direct-drive, gas turbine engine 100 for an aircraft, that mayincorporate one or more embodiments of the present disclosure. The gasturbine engine 100 includes a fan assembly 102 (e.g., a fixed-pitch fanassembly) and a turbomachine 103 (also referred to as a core engine). Inthis example, the turbomachine 103 is a two-spool turbomachine, whichhas a high-speed system and a low-speed system.

The high-speed system of the turbomachine 103, which is not shown inFIG. 1 , includes a high-pressure compressor, a high-pressure turbine, acombustor, and a high-speed shaft (also referred to as a “high-pressureshaft”) supported by bearings and connecting the high-pressurecompressor and the high-pressure turbine. The high-speed shaft,components of the high-pressure compressor, and components of thehigh-pressure turbine all rotate around a centerline axis 112 of the gasturbine engine 100. The high-pressure compressor (or at least therotating components thereof), the high-pressure turbine (or at least therotating components thereof), and the high-speed shaft may becollectively referred to as a high-pressure spool of the gas turbineengine 100. The combustor is located between the high-pressurecompressor and the high-pressure turbine. The combustor receives amixture of fuel and air, and provides a flow of combustion gases throughthe high-pressure turbine for driving the high-pressure spool. Thehigh-pressure compressor, the high-pressure turbine, and the combustortogether define an engine core of the turbomachine 103.

The low-speed system of the turbomachine 103 includes a low-pressureturbine 120, a low-pressure compressor or booster 121, and a low-speedshaft 122 (also referred to as a “low-pressure shaft”) extending betweenand connecting the booster 121 and the low-pressure turbine 120. In someembodiments, the low-speed shaft 122 may extend further along thecenterline axis 112 than is shown in FIG. 1 . The low-pressure turbine120 is sometimes referred to as the engine's power turbine. Thelow-pressure turbine 120 converts kinetic energy contained in the hotgas exiting from the high-pressure turbine into mechanical shaft energy(e.g., of the low-speed shaft 122), which drives the booster 121 and thefan blades 124 either directly or through a gearbox (e.g., any of thegearboxes detailed herein).

As shown in FIG. 1 , the gas turbine engine 100 defines an axialdirection A (extending parallel to the centerline axis 112), a radialdirection R that extends outward from, and inward to, the centerlineaxis 112 in a direction orthogonal to the axial direction A, and acircumferential direction C that extends three hundred sixty degrees(360°) around the centerline axis 112.

The low-speed shaft 122 is supported on bearings 123 a, 123 b, 123 c,123 d, which are mounted to support structures (not shown) of the gasturbine engine 100. At each position, only two bearings are shown inFIG. 1 for clarity, though more than two bearings, e.g., 3 or 4 bearingsforward and/or aft of the respective illustrated locations, may bearranged to support the low-speed shaft 122 at the respective positions,and may be evenly spaced or irregularly spaced depending on the geometryof the bearing supporting structure, and available space and clearances.

The low-speed shaft 122, components of the booster 121, and componentsof the low-pressure turbine 120 all rotate around the centerline axis112 of the gas turbine engine 100, in either the same direction or acounter-rotating direction as that of the high-pressure spool. Thebooster 121 (or at least the rotating components thereof), thelow-pressure turbine 120 (or at least the rotating components thereof),and the low-speed shaft 122 may collectively be referred to as alow-pressure spool 400 of the gas turbine engine 100, and is furtherdescribed in FIG. 4 .

The fan assembly 102 includes an array of fan blades 124 extendingradially outward from a rotor disc 126. The rotor disc 126 is covered bya rotatable fan hub 127 aerodynamically contoured to promote an airflowthrough array of fan blades 124. The gas turbine engine 100 has anintake side 128 and an exhaust side 130.

The turbomachine 103 is generally encased in a cowl 131. Moreover, itwill be appreciated that the cowl 131 defines at least in part an inlet132 of the turbomachine 103 and an exhaust nozzle 135 of theturbomachine 103, and includes a turbomachinery flow path extendingbetween the inlet 132 and the exhaust nozzle 135. For the embodimentshown in FIG. 1 , the inlet 132 has an annular or an axisymmetric threehundred sixty-degree configuration, and provides a flow path forincoming atmospheric air to enter the turbomachine 103. Such a locationmay be advantageous for a variety of reasons, including management oficing performance as well as protecting the inlet 132 from variousobjects and materials as may be encountered in operation.

For a ducted turbofan engine (FIG. 1 ), a nacelle 140 or fan ductsurrounds the array of fan blades 124. The nacelle 140 is supportedrelative to the turbomachine 103 by circumferentially spaced outletguide vanes 142. The portion of air entering the nacelle 140 andbypassing the inlet 132 to the core engine is called the bypass airflow.In the embodiment of FIG. 1 , the bypass airflow flows through a bypassairflow passage 146 defined at a downstream end 144 of the nacelle 140.

For reference purposes, FIG. 1 depicts a forward or thrust directionwith arrow F, which in turn defines the forward and aft portions of thesystem. The fan assembly 102 is forward of the turbomachine 103, and theexhaust nozzle 135 is aft. The fan assembly 102 is driven by theturbomachine 103, and, more specifically, is driven by the low-pressureturbine 120.

In operation, a volume of air flows through fan assembly 102, and as thevolume of air passes across the array of fan blades 124, a first portionof air is directed or routed into the bypass airflow passage 146, and asecond portion of air is directed or routed into the inlet 132 and alongthe turbomachinery flow path. The ratio between the volume of the firstportion of air and the volume of the second portion of air is commonlyknown as a bypass ratio.

After entering the inlet 132, the second portion of air enters thebooster 121 and the high-pressure compressor (not shown in FIG. 1 ). Thehighly compressed air proceeds along the turbomachinery flow path and isdelivered to the combustor (not shown in FIG. 1 ), where the compressedair is mixed with fuel and burned to provide combustion exhaust gases.The exhaust from the combustor drives the high-pressure turbine (notshown in FIG. 1 ) and the low-pressure turbine 120, and the low-pressureturbine 120 drives the fan assembly 102 via the low-speed shaft 122.

The combustion exhaust gases are subsequently routed through the exhaustnozzle 135 to provide propulsive thrust. Simultaneously, the pressure ofthe first portion of air is substantially increased as the first portionof air is routed through the bypass airflow passage 146 before beingexhausted from a fan exhaust 148 at the downstream end 144, alsoproviding propulsive thrust. In such a manner, the fan blades 124 of thefan assembly 102 are driven to rotate around the centerline axis 112 andgenerate thrust to propel the gas turbine engine 100, and, hence, anaircraft to which it is mounted, in the forward direction F. Otherconfigurations are possible and contemplated within the scope of thepresent disclosure, such as what may be termed a “pusher” configurationembodiment in which the turbomachine 103 is located forward of the fanassembly 102.

As shown, the gas turbine engine 100 in the embodiment shown in FIG. 1has a direct drive configuration in which the low-speed shaft 122 isdirectly coupled to the rotor disc 126 and thereby rotates the fanassembly 102 at the same rotational speed as the low-pressure spool.Alternatively, in some embodiments, the turbomachine 103 includes apower gearbox (not shown in FIG. 1 ), and the fan assembly 102 isindirectly driven by the low-pressure spool of the turbomachine 103across the power gearbox. The power gearbox may include a gearset fordecreasing a rotational speed of the low-pressure spool relative to thelow-pressure turbine 120, such that the fan assembly 102 may rotate at aslower rotational speed than does the low-pressure spool.

FIG. 2 shows a schematic, cross-sectional view of a ducted,indirect-drive, gas turbine engine 200, also referred to as turbineengine 200, taken along a centerline axis 212 of the gas turbine engine200, according to an embodiment of the present disclosure. The gasturbine engine 200, also referred to herein as a turbine engine 200, issimilar in some respects to the gas turbine engine 100 discussed abovewith respect to FIG. 1 , and like reference numerals have been used torefer to the same or similar components. Parts omitted from FIG. 1 forclarity are shown and described with respect to FIG. 2 and, thus, theparts referenced, but not shown, in FIG. 1 may be the same or similarthose shown and described with respect to FIG. 2 . Likewise, partsomitted from the description of FIG. 2 for clarity are shown anddescribed with respect to FIG. 1 , and thus, the parts depicted but notdescribed may be the same as, or similar to, the parts described withrespect to FIG. 1 .

As shown in FIG. 2 , the turbine engine 200 includes, in downstreamserial flow relationship, a fan section 214 including a fan 202, acompressor section 216 including a booster or low-pressure (LP)compressor 221 and a high-pressure (HP) compressor 218, a combustionsection 228 including a combustor 230, a turbine section 233 includingan HP turbine 234, and an LP turbine 220, and an exhaust section 238.

The fan section 214 includes a fan casing 240, which is secured to anacelle (FIG. 1 ) surrounding the fan 202. The fan 202 includes aplurality of fan blades 224 disposed radially about the centerline axis212. The HP compressor 218, the combustor 230, and the HP turbine 234form an engine core 244 of the turbine engine 200, which generatescombustion gases. The engine core 244 is surrounded by a core casing231, which is coupled to the fan casing 240. The fan casing 240 issupported relative to the turbomachine by circumferentially spacedoutlet guide vanes 282.

A high-speed shaft 248 is disposed coaxially about the centerline axis212 of the turbine engine 200 and drivingly connects the HP turbine 234to the HP compressor 218. A low-speed shaft 222, which is disposedcoaxially about the centerline axis 212 of the turbine engine 200 andwithin the larger diameter annular high-speed shaft 248, drivinglyconnects the LP turbine 220 to the LP compressor 221 and the fan 202(either directly or through a gearbox assembly 250). The high-speedshaft 248 and the low-speed shaft 222 are rotatable about the centerlineaxis 212.

The LP compressor 221 and the HP compressor 218, respectively, include arespective plurality of compressor stages 252, 254, in which arespective set of compressor blades 256, 258 rotate relative to arespective set of compressor vanes 260, 262 to compress or pressurizegas entering through the inlet 232. Referring now only to the HPcompressor 218, a single compressor stage 254 includes multiplecompressor blades 258 provided on a rotor disk 261 (or blades and diskare integrated together, referred to as a blisk). A compressor bladeextends radially outwardly relative to the engine centerline 212, from ablade platform to a blade tip The compressor vanes 262 are positionedupstream/downstream of and adjacent to rotating compressor blades 258.The disk 261 for a stage of compressor blades 258 is mounted to thehigh-speed shaft 248 (HPC). A stage of the HPC refers to a single diskof rotor blades or both the rotor blades and adjacent stator vanes (itis understood that either meaning can apply within the context of thisdisclosure without loss of clarity).

The HP turbine 234 has one or two stages 264. In a single turbine stage264 turbine blades 268 are provided on a rotor disk 271. A turbine bladeextends radially outwardly relative to the centerline axis 212, from ablade platform to a blade tip. The HP turbine 234 can also include astator vane 272. The HP turbine 234 may have both an upstream nozzleadjacent the combustor exit and an exit nozzle aft of the rotor, or anozzle upstream of rotor blades or downstream of the rotor blades.

Air exiting the HP turbine 234 enters the LP turbine or power turbine220, which has a plurality of stages of rotating blades 270. The LPturbine 220 can have three, four, five, or six stages. In a single LPturbine stage 266 (containing a plurality of blades coupled to the LPshaft 222), a turbine blade is provided on a rotor disk (connected tothe LP shaft 222) and extends radially outwardly relative to thecenterline axis 212, from a blade platform to a blade tip. The LPturbine 220 can also include a stator vane 274. The LP turbine 220 mayhave both an upstream nozzle and an exit nozzle aft of a stage, followedby the engine's exhaust nozzle 238.

The turbine engine 200 of FIG. 2 operates in a similar manner as theengine of FIG. 1 . Airflow exiting the fan section 214 is split suchthat a portion of the airflow is channeled into an inlet 232 to the LPcompressor 221, which then supplies pressurized airflow to the HPcompressor 218, which further pressurizes the air. The pressurizedairflow from the HP compressor 218 is mixed with fuel in the combustor230 and ignited, thereby generating combustion gases. Some work isextracted from the combustion gases by the HP turbine 234, which drivesthe HP compressor 218 to produce a self-sustaining combustion. Thecombustion gases discharged from the HP turbine enter the LP turbine220, which extracts additional work to drive the LP compressor 221 andthe fan 202 (directly or through the gearbox assembly 250). The gasdischarged from the LP turbine exits through the exhaust nozzle 238.

Some of the air supplied by the fan 202 bypasses the engine core 244 andis used for cooling of portions, especially hot portions, of the turbineengine 200, and/or used to cool or power other aspects of the aircraft.In the context of the turbine engine 200, the hot portions refer to avariety of portions of the turbine engine 200 downstream of thecombustion section 228 (e.g., the turbine section 233). Other sources ofcooling fluid include, but are not limited to, fluid discharged from theLP compressor 221 or the HP compressor 218.

The gas turbine engines 100 and 200 depicted in FIG. 1 and FIG. 2 are byway of example only. In other embodiments, the gas turbine engine mayhave any other suitable configuration, including, for example, any othersuitable number or configurations of shafts or spools, fan blades,turbines, compressors, or combination thereof. The gearbox assembly mayhave any suitable configuration, including, for example, a star gearconfiguration, a planet gear configuration, a single-stage, amulti-stage, epicyclic, non-epicyclic, etc., as detailed further below.The gearbox may have a gear ratio of in a range of 3:1 to 4:1, 3:5 to4:1, 3.25:1 to 3.5:1, or 4:1 to 5:1. The fan assembly may be anysuitable fixed-pitched assembly or variable-pitched assembly. The gasturbine engine may include additional components not shown in FIG. 1 ,such as rotor blades, stator vanes, etc. The fan assembly may beconfigured in any other suitable manner (e.g., as a fixed pitch fan) andfurther may be supported using any other suitable fan frameconfiguration. Aspects of the present disclosure may be incorporatedinto any other suitable turbine engine, including, but not limited to,turbofan engines, propfan engines, turbojet engines, turboprop, andturboshaft engines.

FIG. 3 shows a schematic view of an unducted, three-stream, gas turbineengine 310 for an aircraft, that may incorporate one or more embodimentsof the present disclosure. The gas turbine engine 310 is a “three-streamengine” in that its architecture provides three distinct streams(labeled S1, S2, and S3) of thrust-producing airflow during operation,as detailed further below.

As shown in FIG. 3 , the gas turbine engine 310 defines an axialdirection A, a radial direction R, and a circumferential direction C.Moreover, the gas turbine engine 310 defines a centerline axis 312 thatextends along the axial direction A. In general, the axial direction Aextends parallel to the centerline axis 312, the radial direction Rextends outward from, and inward to, the centerline axis 312 in adirection orthogonal to the axial direction A, and the circumferentialdirection C extends three hundred sixty degrees (360°) around thecenterline axis 312. The gas turbine engine 310 extends between aforward end 314 and an aft end 316, e.g., along the axial direction A.

The gas turbine engine 310 includes a core engine 320 and a fan assembly350 positioned upstream thereof. Generally, the core engine 320includes, in serial flow order, a compressor section, a combustionsection, a turbine section, and an exhaust section. Particularly, asshown in FIG. 3 , the core engine 320 includes an engine core 318 and acore cowl 322 that annularly surrounds the core engine 320. The coreengine 320 and core cowl 322 define a core inlet 324 having an annularshape. The core cowl 322 further encloses and supports a low pressure(LP) compressor 326 (also referred to as a booster) for pressurizing theair that enters the core engine 320 through core inlet 324. A highpressure (HP) compressor 328 receives pressurized air from the LPcompressor 326 and further increases the pressure of the air. Thepressurized air flows downstream to a combustor 330 where fuel isinjected into the pressurized air and ignited to raise the temperatureand the energy level of the pressurized air, thereby generatingcombustion gases.

The combustion gases flow from the combustor 330 downstream to a highpressure (HP) turbine 332. The HP turbine 332 drives the HP compressor328 through a first shaft, also referred to as a high pressure (HP)shaft 336 (also referred to as a “high-speed shaft”). In this regard,the HP turbine 332 is drivingly coupled with the HP compressor 328.Together, the HP compressor 328, the combustor 330, and the HP turbine332 define the engine core 318. The combustion gases then flow to apower turbine or low pressure (LP) turbine 334. The LP turbine 334drives the LP compressor 326 and components of the fan assembly 350through a second shaft, also referred to as a low pressure (LP) shaft338 (also referred to as a “low-speed shaft”). In this regard, the LPturbine 334 is drivingly coupled with the LP compressor 326 andcomponents of the fan assembly 350. The low-speed shaft 338 is coaxialwith the high-speed shaft 336 in the embodiment of FIG. 3 . Afterdriving each of the HP turbine 332 and the LP turbine 334, thecombustion gases exit the core engine 320 through a core exhaust nozzle340. The core engine 320 defines a core flowpath, also referred to as acore duct 342, that extends between the core inlet 324 and the coreexhaust nozzle 340. The core duct 342 is an annular duct positionedgenerally inward of the core cowl 322 along the radial direction R.

The fan assembly 350 includes a primary fan 352. For the embodiment ofFIG. 3 , the primary fan 352 is an open rotor fan, also referred to asan unducted fan. However, in other embodiments, the primary fan 352 maybe ducted, e.g., by a fan casing or a nacelle circumferentiallysurrounding the primary fan 352. The primary fan 352 includes an arrayof fan blades 354 (only one shown in FIG. 3 ). The fan blades 354 arerotatable about the centerline axis 312 via a fan shaft 356. As shown inFIG. 3 , the fan shaft 356 is coupled with the low-speed shaft 338 via aspeed reduction gearbox, also referred to as a gearbox assembly 355,e.g., in an indirect-drive configuration. The gearbox assembly 355 isshown schematically in FIG. 3 . The gearbox assembly 355 includes aplurality of gears for adjusting the rotational speed of the fan shaft356 and, thus, the primary fan 352 relative to the low-speed shaft 338to a more efficient rotational fan speed. The gearbox assembly may havea gear ratio of 4:1 to 12:1, or 7:1 to 12:1, or 4:1 to 10:1, or 5:1 to9:1, or 6:1 to 9:1, and may be configured in an epicyclic star or planetgear configuration. The gearbox may be a single stage or compoundgearbox.

The fan blades 354 can be arranged in equal spacing around thecenterline axis 312. Each fan blade 354 has a root and a tip and a spandefined therebetween. Each fan blade 354 defines a central blade axis357. For the embodiment of FIG. 3 , each fan blade 354 of the primaryfan 352 is rotatable about their respective central blade axis 357,e.g., in unison with one another. One or more actuators 358 arecontrolled to pitch the fan blades 354 about their respective centralblade axis 357. In other embodiments, each fan blade 354 is fixed or isunable to be pitched about the central blade axis 357.

The fan assembly 350 further includes a fan guide vane array 360 thatincludes fan guide vanes 362 (only one shown in FIG. 3 ) disposed aroundthe centerline axis 312. For the embodiment of FIG. 3 , the fan guidevanes 362 are not rotatable about the centerline axis 312. Each fanguide vane 362 has a root and a tip and a span defined therebetween. Thefan guide vanes 362 can be unshrouded as shown in FIG. 3 or can beshrouded, e.g., by an annular shroud spaced outward from the tips of thefan guide vanes 362 along the radial direction R. Each fan guide vane362 defines a central vane axis 364. For the embodiment of FIG. 3 , eachfan guide vane 362 of the fan guide vane array 360 is rotatable abouttheir respective central vane axis 364, e.g., in unison with oneanother. One or more actuators 366 are controlled to pitch the fan guidevanes 362 about their respective central vane axis 364. In otherembodiments, each fan guide vane 362 is fixed or is unable to be pitchedabout the central vane axis 364. The fan guide vanes 362 are mounted toa fan cowl 370.

The fan cowl 370 annularly encases at least a portion of the core cowl322 and is generally positioned outward of the core cowl 322 along theradial direction R. Particularly, a downstream section of the fan cowl370 extends over a forward portion of the core cowl 322 to define a fanflowpath, also referred to as a fan duct 372. Incoming air entersthrough the fan duct 372 through a fan duct inlet 376 and exits througha fan exhaust nozzle 378 to produce propulsive thrust. The fan duct 372is an annular duct positioned generally outward of the core duct 342along the radial direction R. The fan cowl 370 and the core cowl 322 areconnected together and supported by a plurality of struts 374 (only oneshown in FIG. 3 ) that extend substantially radially and arecircumferentially spaced about the centerline axis 312. The plurality ofstruts 374 are each aerodynamically contoured to direct air flowingthereby. Other struts in addition to the plurality of struts 374 can beused to connect and support the fan cowl 370 and/or the core cowl 322.

The gas turbine engine 310 also defines or includes an inlet duct 380.The inlet duct 380 extends between an engine inlet 382 and the coreinlet 324 and the fan duct inlet 376. The engine inlet 382 is definedgenerally at the forward end of the fan cowl 370 and is positionedbetween the primary fan 352 and the fan guide vane array 360 along theaxial direction A. The inlet duct 380 is an annular duct that ispositioned inward of the fan cowl 370 along the radial direction R. Airflowing downstream along the inlet duct 380 is split, not necessarilyevenly, into the core duct 342 and the fan duct 372 by a splitter 384 ofthe core cowl 322. The inlet duct 380 is wider than the core duct 342along the radial direction R. The inlet duct 380 is also wider than thefan duct 372 along the radial direction R.

The fan assembly 350 also includes a mid-fan 386. The mid-fan 386includes a plurality of mid-fan blades 388 (only one shown in FIG. 3 ).The plurality of mid-fan blades 388 are rotatable, e.g., about thecenterline axis 312. The mid-fan 386 is drivingly coupled with the LPturbine 334 via the low-speed shaft 338. The plurality of mid-fan blades388 can be arranged in equal circumferential spacing about thecenterline axis 312. The plurality of mid-fan blades 388 are annularlysurrounded (e.g., ducted) by the fan cowl 370. In this regard, themid-fan 386 is positioned inward of the fan cowl 370 along the radialdirection R. The mid-fan 386 is positioned within the inlet duct 380upstream of both the core duct 342 and the fan duct 372. A ratio of aspan of a blade 354 to that of a mid-fan blade 388 (a span is measuredfrom a root to tip of the respective blade) is greater than 2 and lessthan 10, to achieve the desired benefits of the third stream (S3),particularly the additional thrust it offers to the engine, which canenable a smaller diameter blade 354 (benefits engine installation).

Accordingly, air flowing through the inlet duct 380 flows across theplurality of mid-fan blades 388 and is accelerated downstream thereof.At least a portion of the air accelerated by the mid-fan blades 388flows into the fan duct 372 and is ultimately exhausted through the fanexhaust nozzle 378 to produce propulsive thrust. Also, at least aportion of the air accelerated by the plurality of mid-fan blades 388flows into the core duct 342 and is ultimately exhausted through thecore exhaust nozzle 340 to produce propulsive thrust. Generally, themid-fan 386 is a compression device positioned downstream of the engineinlet 382. The mid-fan 386 is operable to accelerate air into the fanduct 372, also referred to as a secondary bypass passage.

During operation of the gas turbine engine 310, an initial or incomingairflow passes through the fan blades 354 of the primary fan 352 andsplits into a first airflow and a second airflow. The first airflowbypasses the engine inlet 382 and flows generally along the axialdirection A outward of the fan cowl 370 along the radial direction R.The first airflow accelerated by the fan blades 354 passes through thefan guide vanes 362 and continues downstream thereafter to produce aprimary propulsion stream or first thrust stream S1. A majority of thenet thrust produced by the gas turbine engine 310 is produced by thefirst thrust stream S1. The second airflow enters the inlet duct 380through the engine inlet 382.

The second airflow flowing downstream through the inlet duct 380 flowsthrough the plurality of mid-fan blades 388 of the mid-fan 386 and isconsequently compressed. The second airflow flowing downstream of themid-fan blades 388 is split by the splitter 384 located at the forwardend of the core cowl 322. Particularly, a portion of the second airflowflowing downstream of the mid-fan 386 flows into the core duct 342through the core inlet 324. The portion of the second airflow that flowsinto the core duct 342 is progressively compressed by the LP compressor326 and the HP compressor 328 and is ultimately discharged into thecombustion section. The discharged pressurized air stream flowsdownstream to the combustor 330 where fuel is introduced to generatecombustion gases or products.

The combustor 330 defines an annular combustion chamber that isgenerally coaxial with the centerline axis 312. The combustor 330receives pressurized air from the HP compressor 328 via a pressurecompressor discharge outlet. A portion of the pressurized air flows intoa mixer. Fuel is injected by a fuel nozzle (omitted for clarity) to mixwith the pressurized air thereby forming a fuel-air mixture that isprovided to the combustion chamber for combustion. Ignition of thefuel-air mixture is accomplished by one or more igniters (omitted forclarity), and the resulting combustion gases flow along the axialdirection A toward, and into, a first stage turbine nozzle of the HPturbine 332. The first stage turbine nozzle 333 is defined by an annularflow channel that includes a plurality of radially extending,circumferentially spaced nozzle vanes 335 that turn the combustion gasesso that they flow angularly and impinge upon first stage turbine bladesof the HP turbine 332. The combustion gases exit the HP turbine 332 andflow through the LP turbine 334 and exit the core duct 342 through thecore exhaust nozzle 340 to produce a core air stream, also referred toas a second thrust stream S2. As noted above, the HP turbine 332 drivesthe HP compressor 328 via the high-speed shaft 336, and the LP turbine334 drives the LP compressor 326, the primary fan 352, and the mid-fan386 via the low-speed shaft 338.

The other portion of the second airflow flowing downstream of themid-fan 386 is split by the splitter 384 into the fan duct 372. The airenters the fan duct 372 through the fan duct inlet 376. The air flowsgenerally along the axial direction A through the fan duct 372 and isultimately exhausted from the fan duct 372 through the fan exhaustnozzle 378 to produce a third stream, also referred to as a third thruststream S3.

The third thrust stream S3 is a secondary air stream that increasesfluid energy to produce a minority of total propulsion system thrust. Insome embodiments, a pressure ratio of the third stream is higher thanthat of the primary propulsion stream (e.g., a bypass or a propellerdriven propulsion stream). The thrust may be produced through adedicated nozzle or through mixing of the secondary air stream with theprimary propulsion stream or a core air stream, e.g., into a commonnozzle. In certain embodiments, an operating temperature of thesecondary air stream is less than a maximum compressor dischargetemperature for the engine. Furthermore in certain embodiments, aspectsof the third stream (e.g., airstream properties, mixing properties, orexhaust properties), and thereby a percent contribution to total thrust,are passively adjusted during engine operation or can be modifiedpurposefully through use of engine control features (such as fuel flow,electric machine power, variable stators, variable inlet guide vanes,valves, variable exhaust geometry, or fluidic features) to adjust or toimprove overall system performance across a broad range of potentialoperating conditions.

The gas turbine engine 310 depicted in FIG. 3 is by way of example only.In other embodiments, the gas turbine engine 310 may have any othersuitable configuration. For example, in other embodiments, the primaryfan 352 may be configured in any other suitable manner (e.g., as a fixedpitch fan) and further may be supported using any other suitable fanframe configuration. In other embodiments, the primary fan 352 can beducted by a fan casing or a nacelle such that a bypass passage isdefined between the fan casing and the fan cowl 370. Moreover, in otherembodiments, any other suitable number or configuration of compressors,turbines, shafts, or a combination thereof may be provided. In stillother embodiments, aspects of the present disclosure may be incorporatedinto any other suitable turbine engine, such as, for example, turbofanengines, propfan engines, turbojet engines, turboprop, turboshaftengines, and/or turbine engines defining two streams (e.g., a bypassstream and a core air stream).

Further, for the depicted embodiment of FIG. 3 , the gas turbine engine310 includes an electric machine 390 (motor-generator) operably coupledwith a rotating component thereof. In this regard, the gas turbineengine 310 is a hybrid-electric propulsion machine. Particularly, asshown in FIG. 3 , the electric machine 390 is operatively coupled withthe low-speed shaft 338. The electric machine 390 can be mechanicallyconnected to the low-speed shaft 338, either directly, or indirectly,e.g., by way of a gearbox assembly 392 (shown schematically in FIG. 3 ).Further, although in this embodiment the electric machine 390 isoperatively coupled with the low-speed shaft 338 at an aft end of thelow-speed shaft 338, the electric machine 390 can be coupled with thelow-speed shaft 338 at any suitable location or can be coupled to otherrotating components of the gas turbine engine 310, such as thehigh-speed shaft 336 or the low-speed shaft 338. For instance, in someembodiments, the electric machine 390 can be coupled with the low-speedshaft 338 and positioned forward of the mid-fan 386 along the axialdirection. In some embodiments, the engine of FIG. 2 also includes anelectric machine coupled to the LP shaft and located in the engine'stail cone.

In some embodiments, the electric machine 390 can be an electric motoroperable to drive or motor the low-speed shaft 338, e.g., during anengine burst. In other embodiments, the electric machine 390 can be anelectric generator operable to convert mechanical energy into electricalenergy. In this way, electrical power generated by the electric machine390 can be directed to various engine and/or aircraft systems. In someembodiments, the electric machine 390 can be a motor/generator with dualfunctionality. The electric machine 390 includes a rotor 394 and astator 396. The rotor 394 is coupled to the low-speed shaft 338 androtates with rotation of the low-speed shaft 338. In this way, the rotor394 rotates with respect to the stator 396, thereby generatingelectrical power. Although the electric machine 390 has been describedand illustrated in FIG. 3 as having a particular configuration, thepresent disclosure may apply to electric machines having alternativeconfigurations. For instance, the rotor 394 and/or the stator 396 mayhave different configurations or may be arranged in a different mannerthan illustrated in FIG. 3 .

FIG. 4 shows an enlarged view of a portion of the cross-sectional viewof FIG. 1 , that includes the low-pressure spool 400 according to someembodiments of the present disclosure. For example, a portion of thebooster 121 and a portion of the low-pressure turbine 120 are shownmounted to the low-speed shaft 122 of the turbomachine 103, which inthis example is a two-spool turbomachine. Alternatively, the low-speedshaft 122 may be an intermediate shaft in a three-spool turbomachine(not shown). The low-speed shaft 122 is supported by at least bearings123 a to 123 d, which are located at mounting points 405 a, 405 bassociated with a booster 121 location and a low-pressure turbine 120location, respectively, for providing shaft rotational support at theselocations. In the example of FIG. 4 , bearings 123 a, 123 b, 123 c, and123 d are all positioned inside of the mounting points 405 a and 405 b,which is referred to as an inbound bearing layout, or alternativelyreferred to as an overhung configuration for the booster 121 andlow-pressure turbine 120. If the bearings were positioned outside of themounting point 405 b, then that would be referred to as an outboundlayout. The bearings 123 a to 123 d can, however, be positioned at anypoint along the low-speed shaft 122, and may both be inbound, both beoutbound, or one inbound and the other outbound.

The low-speed shaft 122 has a length “L” (indicated by arrow 408) and anouter diameter “D” (indicated by arrow 410). The length L is alsoreferred to as L_(MSR) and the outer diameter D is also referred toD_(MSR), as detailed further below. The low-speed shaft 122 can behollow, with an inner diameter “d” indicated by arrow 411). In caseswhen the diameter of the low-speed shaft 122 varies along the length L,the outer diameter “D” and the inner diameter “d” may be defined at amidpoint of the low-speed shaft 122 (also referred to as the midshaft415). The thickness may be defined as the thickness of the walls of thelow-speed shaft 122 in embodiments in which the low-speed shaft 122 ishollow. A difference between a stated outer diameter D and innerdiameter d of the low-speed shaft 122 may be understood as the shaft'swall thickness. In cases when the wall thickness varies along the lengthof the low-speed shaft 122, the thickness may be defined as thedifference between the inner diameter and the outer diameter at themidshaft 415.

In some embodiments, the length L can be understood as the portion ofthe low-speed shaft 122 between the bearings 123 a to 123 d and/or themounting points 405 a, 405 b of engine components such as the booster121 and the low-pressure turbine 120. For example, in the two-bearingarrangement of FIG. 4 , the length L may be measured as the distancebetween midpoints of the bearings 123 a to 123 b, as indicated by thedashed vertical lines and arrow 408. For a four-bearing arrangement,there may be additional bearings along the shaft, in which case thelength L may be measured as the distance between the midpoints of aninnermost pair of bearings, or the distance between pairs or othergroupings of bearings. In some embodiments, the length may be measuredrelative to specific bearings associated with specific engine componentssuch as the booster 121 and the low-pressure turbine 120.

During operation, the low-speed shaft 122 rotates with a rotationalspeed that can be expressed in either rotations per minute (RPM), or asan outer diameter (OD) speed expressed in units of linear velocity, suchas feet per second (ft/sec). The rotational stability of the low-speedshaft 122 relative to its operational range may be characterized by theresonance frequency of the fundamental or first order bending mode. Whenan operational speed is the same as this resonance frequency, the shaftis operating at its critical speed. The low-speed shaft 122, whensupported by bearings 123 a to 123 d, has a mode shape for thisfirst-order bending mode that may be generally described as ahalf-sinusoid, with a midshaft 415 location undergoing maximumdisplacement (indicated by arrow 420, which is exaggerated for clarityand is not to scale) and, therefore, having a maximum kinetic energy ofdisplacement relative to other portions of the low-speed shaft 122. Thefundamental mode shape is illustrated by dashed line 425 extending frombearing 123 c to bearing 123 d in FIG. 4 , though this is only half ofthe amplitude of oscillation. This unstable mode is a standing waveacross the length L of the low-speed shaft 122. The maximum deflectionoccurs when the excitation source has a periodicity or cyclic componentnear to the fundamental frequency. Since the bending mode is not activeat the location of the innermost bearings 123 a to 123 d for thelow-speed shaft 122, the instability cannot be mitigated with the use ofbearing dampers. When an engine is designed, the shaft speed expected toproduce the highest deflection or instability at the midshaft is theshaft speed that equals the critical speed.

If the critical speed of the shaft critical speed falls within thestandard operational range, i.e., if the critical speed is below theredline speed or the low-speed shaft 122 is a supercritical shaft, thenduring routine operation, the low-speed shaft 122 may at times operateat or pass through the critical speed, which induces an unstablecondition. Even if the engine is operated at the critical speedtemporarily, there is a possibility of undetected vibration, whirlinstability, and some likelihood of damage. For low vibration andstability, it is preferable to have an operating range free of anyintervening critical speeds.

There is a desire to pursue engines capable of operating at higherredline speeds. This pursuit of higher operating speeds requires thatthe low-speed shaft 122 have a higher strength to weight characteristicif it is also desired that the shaft remain subcritical. The inventorssought this end result—higher speed shafts while remaining subcritical.To this end, a large number of engine designs were evaluated. Dependingon the architecture, the positions and numbers of bearings relative tomounting points 405 a, 405 b were varied, and the resulting impact, notonly on the critical speed but also the feasibility of suchconfigurations given competing requirements (clearance, spacing, sumplocations, oil supply lines), were taken into consideration, as will bereadily apparent in view of the disclosure. A discussion of theseembodiments follows. In the following discussion, strength to weightratio is represented as E/rho, calculated as the ratio of Young'smodulus E for the material (expressed, for example, in pounds per squarefoot) divided by the density rho (expressed, for example, in pounds percubic inch). The shaft bending mode is represented as the criticalrotational speed expressed in rotations per minute (RPM), though itcould alternatively be expressed as the fundamental frequency of thebending mode in Hertz.

In some embodiments, high strength steel alloys, advanced materials,composite materials, and combinations thereof, were contemplated. Forexample, high strength-to-weight ratio materials such as titanium boride(TiB), a titanium metal matrix composite (TiMMC), provided 30% to 50%increased strength-to-weight ratio relative to steel or titanium alloys.In addition, coatings with materials such as graphene were found toimprove strength by a factor of two in lab tests, without impactingweight. These types of changes in material composition may becharacterized in some embodiments by the ratio of E/rho.

FIG. 5A shows a cross-sectional view of a steel shaft 505, with astandard hollow interior 506 surrounded by a steel layer 507, andgeometry defined by a length L, outer diameter D, inner diameter d, etc.

FIG. 5B shows a cross-sectional view of an example of a composite shaft510, with identical geometry to the steel shaft 505. Rather than beingcomposed entirely of steel, the composite shaft 510 has an inner layer515 surrounding a hollow interior 517, a middle layer 520, and an outercoating 525, all of different materials. The middle layer 520 in thisexample is also steel, though in other embodiments the composite shaftcould use no steel at all, or have a different layer be steel.

For example, both the steel shaft 505 and the composite shaft 510 havelength L of seventy-six inches and outer diameter of three inches, alongwith a standard inbound two-bearings configuration as depicted in FIG. 4. The fundamental frequency of the unstable mode for the steel shaft 505is eighty Hertz (Hz), whereas the fundamental frequency for thecomposite shaft 510 is ninety Hz.

In other embodiments, more layers or fewer layers may be used. Some orall of these layers and coatings may be of numerous alternativematerials to steel, including but not limited to TiB, TiMMC, othermetals and metal matrix composites, silicon carbide (SiC), siliconcarbide reinforced metals or alloys (e.g., SiC-MMC), aluminum alloys,graphene, or combinations thereof. The concepts of the presentdisclosure are not limited by the particular materials used for thelayers and coatings. For the composite shaft 510, the critical speedcorresponding to the unstable mode is increased relative to the(otherwise identical) steel shaft 505, which means that relative to thesteel shaft 505, the composite shaft 510 can attain a higher rotationalspeed before reaching the critical speed.

Depending on the type of composite materials chosen and the relativethickness and arrangement of the layers, the ratio of stiffness toweight can be modified, and, therefore the critical speed can beincreased. The inventors conceived of a variety of embodiments resultingfrom the selection of different composite materials, thicknesses, andbearings configurations to allow for operation at higher speed. Two suchembodiments are listed in TABLE 1. These embodiments were considered aspossible designs that could increase the shaft stiffness to weight ratioin such a way to be compatible with engine architecture and withoutrequiring modifications or limitations on the targeted operating rangefor a subcritical shaft.

TABLE 1 Embodi- L D E/rho Teff Mode ment (in) (in) Bearing type (in⁻¹)(in) (RPM) 1 82.2 2.74 2-bearing outbound 1.00E+08 0.35 4181 2 60.6 2.75inbound OTM 1.27E+08 0.35 10263 3 82.2 2.74 outbound OTM 1.27E+08 0.356915

Embodiment 1 was evaluated using a high strength steel alloy and anoutbound bearing layout. Embodiments 2 and 3 were evaluated with acomposite material instead of steel alloy. Embodiment 2 uses overturningmoment (OTM) bearings with an inbound bearing layout that is differentfrom the layout used by Embodiment 1. Embodiment 3 uses OTM bearingswith an outbound bearing layout that is similar to that used byEmbodiment 1. These bearing types and layouts are described in furtherdetail below with reference to FIG. 7A, FIG. 20B, and TABLE 3. Thevalues shown in TABLE 1 illustrate that Embodiments 2 and 3 achieve ahigher strength-to-weight ratio (E/rho) when using a composite material,instead of the steel alloy used in Embodiment 1. As a result of thesedifferences, the shaft mode critical speed occurs at 4181 RPM forEmbodiment 1, at 10263 RPM for Embodiment 2 and at 6915 for Embodiment3.

The inventors also modified the shaft thickness along its length, toevaluate the effect on critical speed for a strength to weight ratio ofE/rho that is not constant along the length L, and for differentsuitable materials. An example of a shaft with a uniform E/rho along itslength L is shown in FIG. 6A, and examples of shafts having variableE/rho are shown in FIG. 6B and FIG. 6C.

FIG. 6A conceptually shows a cross-sectional view of a uniform shaft 605with a constant diameter and thickness. In this example, the uniformshaft 605 has a length L of seventy-six inches. The outer diameter D ofthe uniform shaft 605 is 3.0 inches. The uniform shaft 605 is hollow,with a constant wall thickness of 0.2 inche and corresponding constantinner radius of 1.3 inches along its length. For this example of auniform shaft 605, and a two-bearing outbound configuration such as inFIG. 4 , the fundamental frequency of the unstable mode is eighty Hz.

FIG. 6B conceptually shows a cross-sectional view of a concave shaft 610with a constant outer diameter D and a variable thickness. Forcomparison, the uniform shaft 605 and the concave shaft 610 have thesame material (e.g., hollow steel), bearings (outbound), and length(seventy-six inches), with a constant outer radius of 1.5 inches alongits length. The outer diameter D of the concave shaft 610 is, therefore,3.0 inches. Unlike the uniform shaft 605, however, the concave shaft 610has a wall thickness of 0.3 inch at the ends 612, 614 (e.g., at thebearings, which are not shown in FIGS. 6A to 6C), and a thinner wallthickness of 0.15 inches in the midshaft region 615. This results in aninner radius of 1.35 inches in the midshaft region 615 and a smallerinner radius of 1.2 inches at the ends 612, 614. The concave shaft 610therefore has a reduced mass density in the midshaft region 615. Toachieve the resulting concave profile, various methods may be used tomanufacture the concave shaft 610, such as a bottle boring technique.

FIG. 6C conceptually shows a cross-sectional of a convex shaft 620 witha variable outer diameter D and a variable thickness. For comparison,the uniform shaft 605 and the convex shaft 620 have the same material(e.g., hollow steel), bearings (outbound), and length (seventy-sixinches), with a constant inner radius of 1.2 inches along its length.Unlike the uniform shaft 605, the convex shaft 620 has a wall thicknessof 0.3 inch at the ends 622, 624, and a thinner wall thickness of 0.15inches in the midshaft region 625, just like the concave shaft 610.Unlike the concave shaft 610, the convex shaft 620 has an outer radiusof 1.5 inches at the ends 622, 624, and a smaller outer radius of 1.35inches in the midshaft region 625. The convex shaft 620 also has areduced mass density in the midshaft region 625.

Since the radius (and, therefore, the diameter) are variable over thelength of the convex shaft 620, the diameter D is defined in someembodiments as the diameter at the midshaft region 625, since this hasthe most relevance to the bending mode and undergoes maximum deflection.In the example of the convex shaft 620, the shaft outer diameter D is2.7 inches in the midshaft region 625. In other embodiments, forexample, embodiments when the radius has multiple minima and/or maxima,the diameter D may be defined at any of those minima or maxima. Toachieve the resulting convex profile, various methods may be used tomanufacture the convex shaft 620, such as external machining.

For both convex and concave thickness profiles, as well as types ofvariable thickness profiles, the thickness may be described using aneffective thickness value, Teff. For a uniform shaft the thickness wouldsimply be the difference between the outer diameter and the innerdiameter. When these values are variable over the length of the shaft,the effective thickness can be calculated as the difference between theeffective outer diameter and effective inner diameter. For example, theeffective thickness may be defined at the midshaft in some embodiments.

With variable thickness, in some embodiments the concave shaft 610 andthe convex shaft 620 can have twenty-five to thirty percent less weightthan the uniform shaft 605 in the midshaft region 615 and 625,respectively. Note that the variation in thickness need not becontinuous, for example a stepped change in geometry could also be used.As a result, the fundamental frequency of the unstable mode for both theconcave shaft 610 and the convex shaft 620 is increased to ninety Hz,which is higher than the eighty Hz fundamental frequency for the uniformshaft 605. In other words, the concave shaft 610 and the convex shaft620 can both attain a higher rotational speed than that of uniform shaft605, before reaching subcritical speeds.

The concave shaft 610 and the convex shaft 620 are examples of differentthickness profiles that may be used in some embodiments. Other thicknessprofiles are also contemplated, which reduce or increase the massdensity of the shaft in the midshaft region. The concepts of the currentdisclosure are not limited by the particular thickness profile used.

Depending on the thickness profile, the ratio of stiffness to weight canbe modified to produce significant changes in the critical speed.embodiments are listed in TABLE 2. These embodiments were considered aspossible designs that could modify the effective thickness in such a wayto be compatible with engine architecture and without requiringmodifications or limitations on the targeted operating range for asubcritical shaft.

TABLE 2 Embodi- L D E/rho Teff Mode ment (in) (in) Bearing type (in⁻¹)(in) (RPM) 4 60.6 2.75 inbound OTM 1.00E+08 0.35 9001 5 82.2 2.74outbound OTM 1.00E+08 0.35 6065 6 60.6 2.75 inbound OTM 1.00E+08 0.3210039 7 82.2 2.74 outbound OTM 1.00E+08 0.32 6942

Embodiments 4, 5, 6, and 7 all use a steel alloy material composition.Embodiments 4 and 6 use an inbound bearing layout with OTM bearings, andEmbodiments 5 and 7 use an outbound bearing layout with OTM bearings.Embodiments 4 and 5 are uniform shafts similar to the Example of FIG.6A. Embodiments 6 and 7, however, have a convex thickness profilesimilar to the example of FIG. 6C, having been manufactured with abottle boring manufacturing technique. The values shown in TABLE 2illustrate that Embodiments 6 and 7 achieve a lower effective thicknessTeff due to their convex profile, instead of the uniform profile forEmbodiments 4 and 5. As a result of these differences, the shaft modecritical speed occurs at 9001 RPM for Embodiment 4, and at 10039 RPM forEmbodiment 6. The shaft mode critical speed occurs at 6065 RPM forEmbodiment 5, and at 6942 RPM for Embodiment 7.

The inventors also conceived of a variety of shafts with modifiedbearing configurations. Bearings are used to provide transverse supportto the shaft along its length. Bearings may be ball-type bearings, whichhave a very small contact area with the shaft to provide less friction,or roller-type bearings, which have a large contact area with the shaftto provide increased rigidity and load bearing. Different types ofbearings may be mixed in various bearing layouts. According toadditional embodiments, different bearing layouts were considered, fordifferent combinations of uniform, convex, and concave shafts, orvarying shaft thickness profiles and material composition in order todetermine which combination would work best for a given architecture andneed, as well as taking into account competing engineering requirements.

A variety of combinations of bearing configurations were contemplated,such as embodiments when the number of bearings in duplex and/orstraddling position relative to engine components (e.g., a booster 721or a low-pressure turbine 720) were changed. Either or both of theengine components mounted to the shaft 722 may be straddled or overhung.It was found that these variations can improve the critical speed and/orbe more suitable to accommodate space limitations, lubrication resourcesor other architecture-imposed limitations. The embodiments includedlocating bearings at different inbound or outbound positions relative tothe mounting points 705 a, 705 b.

Specific bearing layouts were preferentially used in variousembodiments. These are now described in more detail, though the conceptsof the present disclosure are not limited by the particular number orarrangement of bearings described herein.

For example, FIG. 7A conceptually shows a low-pressure turbine 720 and abooster 721 mounted on a shaft 722 (e.g., a low-speed shaft) supportedby a four-bearing straddle configuration. Additional bearings locatedaround the circumference of the shaft 722 are omitted from FIG. 7A forclarity. In this system, one pair of bearings 723 a, 724 a straddle(i.e., placed forward and aft of) a mounting point 705 a of the booster721, and a second pair of bearings 723 b, 724 b straddle a mountingpoint 705 b of the low-pressure turbine 720. In this example, bearings724 a, 723 b, and 724 b are roller bearings, and bearing 723 a is a ballbearing, though these bearing types may vary in other embodiments. Thelength L for shaft 722 is represented in some embodiments as thedistance between the midpoints or centers of the innermost bearings 724a, 723 b. The four-bearing straddle layout is used in severalembodiments described with reference to TABLE 3.

As another example, FIG. 7B conceptually shows a low-pressure turbine720 and a booster 721 mounted on a shaft 722 supported by a four-bearingoutbound configuration. Additional bearings located around thecircumference of the shaft 722 are omitted from FIG. 7B for clarity.This system is similar to that of the straddle system shown in FIG. 7A,but differs in that bearings 723 a, 724 a are both placed forward ofmounting point 705 a of the booster 721, and bearings 723 b, 724 b areplaced aft of mounting point 705 b of the low-pressure turbine 720. Theshaft 722 may extend beyond bearings 723 b, 724 b. As in the example ofFIG. 7A, bearings 724 a, 723 b, and 724 b are roller bearings, andbearing 723 a is a ball bearing, though these bearing types may vary.The length L for shaft 722 is represented in some embodiments as thedistance between the midpoints or centers of the innermost bearings 724a, 723 b.

As yet another example, FIG. 7C conceptually shows a shaft 722 with aninbound duplex bearing configuration. Additional bearings located aroundthe circumference of the shaft 722 are omitted from FIG. 7C for clarity.According to some embodiments, a first pair of ball bearings 725 a, 726a is arranged in a duplex configuration aft of the mounting point 705 afor the booster 721. A second pair of ball bearings 725 b, 726 b isarranged in a duplex configuration forward of the mounting point 705 bfor the low-pressure turbine 720. Duplex bearing arrangements may alsobe referred to as double-row bearings, or overturning moment (OTM)bearings, since they provide moment stiffness to the shaft, i.e.,provide resistance to rotation across the bearing locations. In someembodiments duplex bearing types may include tandem bearings,back-to-back bearings, face-to-face bearings, and/or tapered rollerbearings.

In the example shown in FIG. 7C, both the first pair of ball bearings725 a, 726 a and the second pair of ball bearings 725 b, 726 b are in aninbound position, i.e., located closer towards the midshaft 727 than therespective mounting points 705 a, 705 b. In this position, the booster721 and the low-pressure turbine 720 are referred to as overhung. Thisinbound OTM layout is used in Embodiments 2, 4, and 6, for example,described above with reference to TABLES 1 and 2.

Alternatively, the first pair of ball bearings 725 a, 726 a and/or thesecond pair of ball bearings 725 b, 726 b may be in an outboundposition, as shown in FIG. 7D, i.e., located farther from the midshaft727 than the respective mounting points 705 a, 705 b of the booster 721and the low-pressure turbine 720. The length L for the duplex bearingconfigurations shown in FIG. 7C and FIG. 7D may be represented in someembodiments as the distance between the midpoints or centers of theinnermost ball bearings 726 a and 725 b, or alternatively, as thedistance between the center of the first pair of ball bearings 725 a,726 a and the center of the second pair of ball bearings 725 b, 726 b.The outbound OTM layout is used in Embodiments 3, 5, and 7, for example,described above with reference to TABLES 1 and 2.

As a further example, FIG. 7E conceptually shows a shaft 722 with atwo-bearing configuration. This configuration employs a first bearing724 a positioned aft of the mounting point 705 a for the booster 721,and a second bearing 724 b positioned aft of the mounting point 705 bfor the low-pressure turbine 720. The length L for this two-bearingconfiguration is represented in some embodiments as the distance betweenthe midpoints or centers of the bearings 724 a, 724 b. Alternativetwo-bearing configurations may position the two bearings in either anoutbound configuration or an inbound configuration. An example of atwo-bearing layout in an inbound configuration is shown in FIG. 1 , thatshows bearings 123 a, 123 b, 123 c, and 123 d are all located inbound ofthe mounting points for the booster 121 and the low-pressure turbine120. Note that in this context, two is the number of bearings along theshaft 722, and does not include additional bearings along thecircumference of the shaft 722. Embodiment 1, described above withreference to TABLE 1, uses a two-bearing layout in an outboundconfiguration (not shown).

In FIGS. 7A to 7E, the lines connecting the booster 721 to mountingpoint 705 a and low-pressure turbine 720 to mounting point 705 b areintended only to indicate schematically the general location of a netforce of the core engine components (e.g., booster 721 or low-pressureturbine 720) acting on the shaft 722 relative to the bearings, and isillustrated in this fashion only for purposes of illustrating arelationship between the nearest engine component relative to thebearing(s). It will be understood that the actual loading on a shaft isdistributed and comes from not only the engine components represented bybooster 721 and low-pressure turbine 720, but other nearby structures aswell. In these embodiments, the primary loading for purposes of thisdisclosure may, however, be thought of simply in terms of enginecomponents attached to the shaft 722 (e.g., low-pressure turbine 720 andbooster 721). It will be understood that the representation shown inFIGS. 7A to 7E is sufficient in defining the parts of the turbomachinethat mostly influence the shaft 722 behavior.

As discussed, at least one bearing may have an overturning moment (OTM)capability, which can resist relative rotation across the bearing in atleast a lateral plane or a vertical plane. These relative rotations mayoccur during bending of the shaft. The position along the shaft of suchbearings with OTM capabilities may directly affect the critical speed,by providing constraints to the relative rotations of the shaft, inaddition to the transverse support function of the bearings.

Examples of embodiments with different bearing arrangements aresummarized in TABLE 3. Generally, the inventors found that the number ofbearings, the position of the bearings and the OTM capability of thebearings can be selected to make a full range of operations subcriticalfor an engine. In other words, the selection of bearing layout canaffect (either increase or decrease) the shaft's critical speed.

TABLE 3 Embodi- L D E/rho Teff Mode ment (in) (in) Bearing type (in⁻¹)(in) (RPM)  8 60.6 2.75 4-bearing straddle 1.00E+08 0.35 7746  9 60.62.75 4-bearing straddle 1.00E+08 0.32 8555 10 60.6 2.75 4-bearingstraddle 1.27E+08 0.35 8832 11 82.2 2.74 4-bearing straddle 1.27E+080.32 9703 12 60.6 2.75 inbound OTM 1.27E+08 0.32 11386 13 82.2 2.74outbound OTM 1.27E+08 0.32 7873

Embodiments 8, 9, 10, and 11 use a four-bearing straddle layout.Embodiments 8 and 9 use steel alloy, while Embodiments 10 and 11 usecomposite materials. Embodiments 8 and 10 have a uniform thicknessprofile, while Embodiments 9 and 11 have a concave thickness profile,manufactured using a bottle boring method. As a result of thesedifferences, the shaft mode critical speed occurs at 7746 RPM forEmbodiment 8, 8555 RPM for Embodiment 9, 8832 RPM for Embodiment 10, and9703 RPM for Embodiment 11.

Embodiments 11, 12, and 13 all use composite material and concavethickness profile via bottle boring. However, Embodiment 11 uses afour-bearing straddle layout, Embodiment 12 uses an inbound OTM bearinglayout, and Embodiment 13 uses an outbound OTM bearing layout. As aresult of these differences, the shaft mode critical speed occurs at9703 RPM for Embodiment 11, 11386 RPM for Embodiment 12, and 7873 RPMfor Embodiment 13.

Embodiment 11 can also be compared to Embodiments 8, 9, and 10 asdescribed with reference to TABLE 3. This allows a comparison of theimpact on critical speed of using composite material, variable thicknessprofile, and both, on a shaft with a four-bearing straddle layout.

Embodiment 12 can be compared to Embodiments 2, 4, and 6 described withreference to TABLE 2. This allows a comparison of the impact on criticalspeed of using composite material, variable thickness profile, and both,on a shaft with an inbound OTM layout.

Embodiment 13 can be compared to Embodiments 3, 5, and 7 described withreference to TABLE 2. This allows a comparison of the impact on criticalspeed of using composite material, variable thickness profile, and both,on a shaft with an outbound OTM layout.

Additionally, Embodiments 2 and 3 (in TABLE 1), and 10 (in TABLE 3) canbe compared, to evaluate the impact on critical speed of using differentbearing layouts on shafts using composite material. Embodiments 6 and 7(in TABLE 2) and 9 (in TABLE 3) can be compared, to evaluate the impacton critical speed of using different bearing layouts on shafts usingconcave thickness profiles.

The embodiments of turbomachine engines, and in particular the shaftsassociated with a power turbine described with reference to FIGS. 5A,5B,6A to 6C, 7A, and 7B, were found to provide an improvement in theperformance of a shaft vis-à-vis its operating range. In addition to thementioned embodiments and those provided in TABLES 1 to 3, the types ofimprovements to the critical speed of the shaft when these features werecombined, taking into consideration the various benefits, as well asdown-sides, to selecting a particular configuration for a turbomachinearchitecture.

Examples of a subcritical shaft with a high redline speed include ashaft with a redline speed of, e.g., 70 ft/sec and adapted for a shaftmode of 5293 RPM, a shaft with a redline speed of, e.g., 75 ft/sec andadapted for a shaft mode of 6380 RPM, and a shaft with a redline speedof, e.g., 181 ft/sec and adapted for a shaft mode of 11410 RPM.

As mentioned earlier, the inventors sought to improve upon the operatingspeed of a low-speed shaft. With regard to the speed of the low-speedshaft, consideration was given not simply to those factors affecting thecore size, but also the stiffness-to-weight ratio and the critical speedof the low-speed shaft. In contrast to existing gas turbine enginesrequiring lower speeds, embodiments considered presented challenges indetermining how the low-speed shaft speed could be increased withoutresulting in an unstable bending mode during regular operation.

Further, a selection of power turbine shaft and bearing arrangement, andlocation of those bearings for a turbomachine takes into considerationother factors, some of which can limit the selection of a shaft. Theinventors however realized during the course of making the severalembodiments referred to in the foregoing that there is a particularrange of designs, constraints on feasible designs that provided anunexpected benefit. The interplay between components can make itparticularly difficult to select or develop one component during enginedesign and prototype testing, especially when some components are atdifferent stages of completion. For example, one or more components maybe nearly complete, yet one or more other components may be in aninitial or preliminary phase where only one (or a few) design parametersare known. It is desired to arrive at what is possible at an early stageof design, so that the down selection of candidate optimal designs,given the tradeoffs, become more possible. Heretofore the process hassometimes been more ad hoc, selecting one design or another withoutknowing the impact when a concept is first taken into consideration.

Even taken separately from the integration of a shaft design with therest of an engine, modifying an existing shaft to increase its criticalspeed is challenging, and the impact of the different types ofimprovements and configurations on critical speed is not easilypredictable without empirical experimentation and simulation, which canbe enormously expensive and time-consuming. In some cases, amodification may even result in lowering the critical speed.

It is desirable to narrow the range of configurations or combination offeatures that can yield favorable results given the constraints of thedesign, feasibility, manufacturing, certification requirements, etc.early in the design selection process to avoid wasted time and effort.During the course of the evaluation of different embodiments as setforth above, the inventors discovered, unexpectedly, that there exists arelationship between the critical speed of the shaft and the ratio L/D,which uniquely identifies a finite and readily ascertainable (in view ofthis disclosure) number of embodiments suitable for a particulararchitecture that can avoid a supercritical or critical shaft situationduring normal operation of an engine. This relationship is referred toby the inventors as the midshaft rating (MSR), and is calculatedaccording to the following relationship (1) between length, diameter anda redline speed (ft/sec) measured at the outer diameter of the shaft:Midshaft Rating MSR=(L _(MSR) /D _(MSR)) (Shaft OD Speed atredline)^(1/2)  (1)

L_(MSR)/D_(MSR) is shaft length divided by effective shaft outerdiameter, L_(MSR)/D_(MSR). The ratio L_(MSR)/D_(MSR) is multiplied withthe square root of the outer diameter (OD) rotation speed (OD Speed) atthe redline speed for the engine architecture. Generally, the lengthL_(MSR) and diameter D_(MSR) are expressed in inches, and the shaft ODredline speed is the linear speed of the shaft surface. The OD redlinespeed in feet per second is calculated as the shaft mode speed (in RPM)multiplied by the outer circumference of the shaft (the outer diameterof the shaft multiplied by pi), and with additional corrections toconvert from inches to feet and from minutes to seconds. Accordingly,the midshaft rating has units of (velocity)^(1/2).

The midshaft rating identifies embodiments for a turbomachine's powerturbine that allow subcritical operation of the engine for a ratedredline speed. TABLE 4 lists embodiments of the turbine shaft along withits associated MSR value. The embodiments can inform one of thedimensions or qualities of the shaft that are believed reasonable andpractical for a shaft according to its basic features and the intended,rated critical speed. In other words, the midshaft rating, and,optionally, the L_(MSR)/D_(MSR) ratio and/or the OD speed at redline,indicates the operating ranges of interest, taking into account theconstraints within which a turbomachine operates, e.g., size,dimensions, cost, mission requirements, airframe type, etc.

In other embodiments, the midshaft rating may also, or alternatively, beused to define the propulsive system operating at a relatively highredline speed. Such things as the requirements of a propulsive system,the requirements of its subsystem(s), airframe integration needs andlimitations, and performance capabilities may, therefore, be summarizedor defined by the midshaft rating.

In still other embodiments, the midshaft rating may additionally providea particularly useful indication of the efficiency and effectiveness ofthe engine during initial development, e.g., as a tool to accept orreject a particular configuration. Thus, the midshaft rating can beused, for example, to guide low-speed shaft development. Therefore, themidshaft rating can also improve the process of developing aturbomachine engine.

Table 4 lists the bearing layout, the strength-to-weight ratio E/rho ininches⁻¹, the effective thickness Teff in inches, the critical speedcorresponding to the shaft's fundamental mode in RPM, the OD linearspeed at redline in ft/sec, the length-to-diameter ratio L/D, alsoreferred to as L_(MSR)/D_(MSR) (dimensionless), and MSR in(ft/sec)^(1/2) for all the embodiments (1 to 13) of Tables 1 to 13, aswell as a number of additional embodiments (14 to 32). As noted above,L/D represents the ratio of the length over the outer diameter. When theshaft has a variable diameter over its length, the outer diameter may bethe diameter at the midshaft. E/rho represents the material compositionof the shaft, and Teff represents an effective wall thickness of theshaft. For shafts with variable thickness over their length, the wallthickness may be the thickness at the midshaft.

TABLE 4 Em- OD bodi- Bearing E/rho Teff Mode Speed MSR ment Layout in⁻¹in RPM L/D ft/sec (ft/sec)^(1/2)  1 2-bearing 1.00E+8 0.35 4181 30 50214 outbound  2 inbound OTM 1.27E+8 0.35 10263 22 123 247  3 outboundOTM 1.27E+8 0.35 691 30 83 275  4 inbound OTM 1.00E+8 0.35 9001 22 108231  5 outbound OTM 1.00E+8 0.35 6065 30 73 257  6 inbound OTM 1.00E+80.32 10039 22 121 242  7 outbound OTM 1.00E+8 0.32 6942 30 83 272  84-bearing 1.00E+8 0.35 7746 22 93 214 straddle  9 4-bearing 1.00E+8 0.328555 22 103 223 straddle 10 4-bearing 1.27E+8 0.35 8832 22 106 229straddle 11 4-bearing 1.27E+8 0.32 9703 30 116 322 straddle 12 inboundOTM 1.27E+8 0.32 11386 22 137 257 13 outbound OTM 1.27E+8 0.32 7873 3094 290 14 4-bearing 1.00E+8 0.35 6262 26 72 219 outbound 15 2-bearingaft 1.27E+8 0.29 8255 21 109 215 16 2-bearing aft 1.27E+8 0.31 13323 14233 216 17 2-bearing aft 1.27E+8 0.47 5667 23 83 210 18 2-bearing aft1.27E+8 0.29 6380 24 83 215 19 2-bearing aft 1.27E+8 0.31 9821 17 154216 20 2-bearing aft 1.27E+8 0.47 4586 26 67 211 21 2-bearing aft1.00E+8 0.23 6380 24 84 217 22 2-bearing aft 1.00E+8 0.25 13493 14 235218 23 2-bearing aft 1.00E+8 0.38 4586 27 62 210 24 2-bearing aft1.27E+8 0.29 6619 25 87 231 25 2-bearing aft 1.27E+8 0.31 11065 17 176232 26 2-bearing aft 1.27E+8 0.47 4852 28 64 224 27 4-bearing 1.00E+80.29 6380 28 75 245 straddle 28 inbound OTM 1.00E+8 0.31 10666 19 165247 29 outbound OTM 1.00E+8 0.47 4586 31 59 239 30 4-bearing 1.27E+80.23 6380 35 70 289 straddle 31 inbound OTM 1.27E+8 0.25 11410 22 181294 32 outbound OTM 1.27E+8 0.38 5293 33 70 276

Embodiments 15 to 26 use a two-bearing aft layout. These embodimentsdiffer in using composite materials, different shaft geometries, andvariable thickness profiles.

Embodiments 15 to 17 use a composite material instead of steel alloy.These embodiments differ in shaft geometry, with different L/D ratiosranging from 14 to 23.

Embodiments 18 to 20 use a material composite instead of a steel alloy.These embodiments also differ from each other in shaft geometry (e.g.,L/D ratio). These also differ from Embodiments 15 to 17, in being longerand thinner, resulting in a higher range of L/D ratio, from 17 to 26.

Embodiments 21 to 23 use a steel alloy, vary the shaft geometry (lengthand/or diameter), and have a concave thickness profile. These differfrom each other in terms of their effective thickness. These embodimentsmay be compared to Embodiments 24 to 26, which use composite materials,vary the shaft geometry (length and/or diameter), and have a concavethickness profile.

Embodiments 27 to 32 use different bearing layouts. Embodiments 27 to 29use steel alloy and have varying geometry. Embodiments 30 to 32 usematerial composites and a concave thickness profile, in addition tovarying geometry.

Based on the experimentation described above, the inventors identifiedembodiments with MSR between two hundred and three hundred (ft/sec)⁻¹and OD redline speeds ranging from fifty to two hundred fifty ft/sec andwith L/D ratio ranging from twelve to thirty-seven were possible andindicated noticeable improvements in subcritical range when the powerturbine shaft incorporates the various aspects of the disclosure.

Table 5 summarizes examples of different operating ranges forembodiments, such as the embodiments listed in Table 4. For example, anembodiment can be configured with a L/D ranging between twelve andtwenty may have an OD speed between one hundred and fifty and twohundred and fifty ft/sec, and a corresponding range of MSR between onehundred ninety and two hundred forty-five (ft/sec)^(1/2). As anotherexample, an embodiment can be configured with a L/D ranging betweensixteen and thirty may have an OD speed between seventy-five and onehundred seventy-five ft/sec, and a corresponding range of MSR betweentwo hundred twelve and two hundred sixty (ft/sec)^(1/2). As stillanother example, an embodiment can be configured with a L/D rangingbetween twenty-six and thirty-seven may have an OD speed between sixtyand ninety ft/sec, and a corresponding range of MSR between two hundredforty-seven and two hundred eighty-seven ft/sec)^(1/2). These low,nominal, and high ranges as summarized in Table 5 are general examples,and individual embodiments may exceed these values.

TABLE 5 Example Limits L/D OD Speed MSR and Ranges (in/in) (ft/sec)(ft/sec)^(1/2) Low limit 12 250 190 20 150 245 Nominal limit 16 175 21230 75 260 High Limit 26 90 247 37 60 287

In view of the foregoing objectives, in at least certain embodiments, apropulsion system is configured to define an MSR greater than onehundred ninety (ft/sec)^(1/2), such as greater than two hundred(ft/sec)^(1/2), such as at least two hundred thirty-five (ft/sec)^(1/2),up to at least three hundred (ft/sec)^(1/2).

In view of the foregoing objectives, in at least certain embodiments, apropulsion system is configured to define an L/D ratio greater thantwelve, such as greater than sixteen, such as at least twenty-six, up toat least thirty-seven.

In view of the foregoing objectives, in at least certain embodiments, apropulsion system is configured to define an OD redline speed greaterthan sixty ft/sec, such as greater than seventy five ft/sec, such as atleast one hundred and fifty ft/sec, up to at least two hundred and fiftyft/sec.

Based on the teachings in this disclosure, and without limiting thedisclosure to only those embodiments explicitly shown, it will beunderstood how both the manner and the degree to which a modification ofshaft length, diameter, material composition, bearings configuration,and thickness profile affects the MSR, and, additionally, the competingrequirements, or requirements for a turbomachine architecture (e.g.,available spacing/packaging, clearance, sump location, lubrication,etc.) for a given MSR.

The gearbox assembly (e.g., any of the gearbox assemblies detailedherein) also affects the MSR. For example, vibrations of the gearboxassembly can excite the low-speed shaft (e.g., at the first-orderbending mode) such that the vibrations of the gearbox assembly cause thelow-speed shaft to vibrate even when the low-speed shaft is operatingsubcritical. Similarly, vibrations of the low-speed shaft can excite thegearbox assembly such that the vibrations of the low-speed shaft causethe gearbox assembly to vibrate, thereby potentially straining themountings of the gearbox assembly beyond a designed limit. In this way,changes in the gearbox assembly affect a vibratory response of thegearbox assembly through the couplings of the gearbox assembly and theengine static structure and/or through the couplings of the gearboxassembly and the low-speed shaft, or vice-versa. The present disclosuredetails how both the manner and the degree to which a modification ofthe gearbox assembly affects the MSR, given the competing requirementsfor a turbomachine architecture (e.g., available spacing/packaging,clearance, sump location, lubrication, etc.) for a given MSR.

FIGS. 8A and 8B illustrate enlarged, schematic cross-sectional views ofa gearbox assembly 838 with a mounting assembly 800 for a gas turbineengine 810, taken at a centerline axis 812 of the gas turbine engine810. A low-speed shaft 836 is shown schematically in FIG. 8A and extendsbeyond the cut view of FIG. 8A to illustrate the length L_(MSR) of thelow-speed shaft 836. The length of the low-speed shaft 836 is detailedabove. The gearbox assembly 838 can be utilized as any of the gearboxassemblies and the gas turbine engine 810 can be any of the gas turbineengines detailed herein. The mounting assembly 800 shown is that for astar configuration gearbox, described in more detail to follow. Thegearbox assembly 838 includes a sun gear 840, a plurality of planetgears 842, and a ring gear 844. A low-pressure turbine (e.g., any of thelow-pressure turbines detailed herein) drives the low-speed shaft 836,which is coupled to the sun gear 840 of the gearbox assembly 838. Thesun gear 840 of the gearbox assembly 838 is coupled via a flex coupling845 to the rotating low-speed shaft 836.

Radially outwardly of the sun gear 840, and intermeshing therewith, isthe plurality of planet gears 842 that are coupled together by a planetcarrier 846. The planet carrier 846 of the gearbox assembly 838 iscoupled, via a flex mount 847, to an engine static structure 819. Theplanet carrier 846 constrains the plurality of planet gears 842 whileallowing each planet gear of the plurality of planet gears 842 to rotateabout its own axis. Radially outwardly of the plurality of planet gears842, and intermeshing therewith, is the ring gear 844, which is anannular ring gear 844. The ring gear 844 is coupled via a fan shaft 848to a fan (e.g., any of the fans or fan assemblies detailed herein) inorder to drive rotation of the fan about the centerline axis 812 of thegas turbine engine 810. The fan shaft 848 is coupled to a fan frame 849via a fan bearing 850. The fan frame 849 couples the rotating ring gear844 of the gearbox assembly 838 and, thus, the rotating fan shaft 848,to the engine static structure 819. The flex coupling 845, the flexmount 847, and the fan frame 849 define the mounting assembly 800 forthe gearbox assembly 838. As described herein, the flex coupling 845,the flex mount 847, and the fan frame 849 may be referred to as mountingmembers.

Although not depicted in FIGS. 8A and 8B for clarity, each of the sungear 840, the plurality of planet gears 842, and the ring gear 844includes teeth about their periphery to intermesh with the other gears.In the example of FIGS. 8A and 8B, the gearbox assembly 838 is a starconfiguration. That is, the ring gear 844 rotates, while the planetcarrier 846 is fixed and stationary. The planet carrier 846 constrainsthe plurality of planet gears 842 such that the plurality of planetgears 842 do not together rotate around the sun gear 840, while alsoenabling each planet gear of the plurality of planet gears 842 to rotateabout its own axis. That is, since the plurality of planet gears 842mesh with both the rotating ring gear 844 as well as the rotating sungear 840, each of the plurality of planet gears 842 rotate about theirown axes to drive the ring gear 844 to rotate about the centerline axis812 due to the rotation of the sun gear 840. The rotation of the ringgear is 844 conveyed to the fan through the fan shaft 848.

FIG. 8B illustrates the mounting assembly 800 of FIG. 8A translated intoa representative vibratory system where each of the flex coupling 845,the flex mount 847, and the fan frame 849 are shown by representativestructural properties of the members, the representative structuralproperties being the structural stiffness (K) and the damping (C) of therespective members of the mounting assembly 800. As shown, each of theflex coupling 845, the flex mount 847, and the fan frame 849 includesthe representative structural properties (structural stiffness anddamping) in each of the lateral direction, the bending direction, andthe torsional direction.

For example, FIG. 8B represents the gearbox supporting structure interms of structural properties characterizing the nature of the couplingbetween the gearbox assembly 838 and the flex coupling 845. The flexcoupling 845 may be represented in terms of a flex coupling lateralstiffness K_(fc) ^(L) a flex coupling bending stiffness K_(fc) ^(B), aflex coupling torsional stiffness K_(fc) ^(T), a flex coupling lateraldamping C_(fc) ^(L), a flex coupling bending damping C_(fc) ^(B), and aflex coupling torsional damping C_(fc) ^(T). In this way, the flexcoupling 845 acts as a dashpot damper that mechanically dampensvibrations through the flex coupling 845 from the gearbox assembly 838and/or from the low-speed shaft 836 (e.g., the midshaft).

FIG. 8B represents the gearbox supporting structure in terms ofstructural properties characterizing the nature of the coupling betweenthe gearbox assembly 838 and the flex mount 847. The flex mount 847 maybe represented in terms of a flex mount lateral stiffness K_(fm) ^(L), aflex mount bending stiffness K_(fm) ^(B), a flex mount torsionalstiffness K_(fm) ^(T), a flex mount lateral damping C_(fm) ^(L), a flexmount bending damping C_(fm) ^(B), and a flex mount torsional dampingC_(fm) ^(T). In this way, the flex mount 847 acts as a dashpot damperthat mechanically dampens vibrations through the flex mount 847 from thegearbox assembly 838 to the engine static structure 819.

FIG. 8B represents the gearbox supporting structure in terms ofstructural properties characterizing the nature of the coupling betweenthe gearbox assembly 838 and the fan frame 849. The fan frame 849 may berepresented in terms of a fan frame lateral stiffness K_(ff) ^(L), a fanframe bending stiffness K_(ff) ^(B), a fan frame torsional stiffnessK_(ff) ^(T), a fan frame lateral damping C_(ff) ^(L), a fan framebending damping C_(ff) ^(B), and a fan frame torsional damping C_(ff)^(T). In this way, the fan frame 849 acts as a dashpot damper thatmechanically dampens vibrations through the fan frame 849 from thegearbox assembly 838 to the engine static structure 819.

FIGS. 9A and 9B illustrate enlarged, schematic side cross-sectionalviews of a gearbox assembly 938 with a mounting assembly 900 for a gasturbine engine 910, taken at a centerline axis 912 of the gas turbineengine 910. A low-speed shaft 936 is shown schematically in FIG. 9A andextends beyond the cut view of FIG. 9A to illustrate the length L_(MSR)of the low-speed shaft 936. The gearbox assembly 938 can be utilized asany of the gearbox assemblies detailed herein and the gas turbine engine910 can be any of the gas turbine engines detailed herein. The mountingassembly 900 shown is that for a planetary configuration gearbox,described in more detail to follow. Similar to the gearbox assembly 838of FIGS. 8A and 8B, the gearbox assembly 938 includes a sun gear 940, aplurality of planet gears 942, and a ring gear 944. A low-pressureturbine drives the low-speed shaft 936, which is coupled to the sun gear940 of the gearbox assembly 938. The sun gear 940 is coupled via a flexcoupling 945 to the low-speed shaft 936. The plurality of planet gears942 are coupled together by a planet carrier 946. In the embodiment ofFIGS. 9A and 9B, the planet carrier 946 is coupled, via a fan shaft 948,to a fan (e.g., any of the fans or fan assemblies detailed herein) todrive rotation of the fan about the centerline axis 912. The fan shaft948 is coupled to a fan frame 949 via a fan bearing 950. The ring gear944 is coupled via a flex mount 947 to an engine static structure 919.The flex coupling 945, the flex mount 947, and the fan frame 949 definethe mounting assembly 900 for the gearbox assembly 938. As describedherein, the flex coupling 945, the flex mount 947, and the fan frame 949may be referred to as mounting members.

In the embodiment of FIGS. 9A and 9B, the gearbox assembly 938 is aplanetary configuration. That is, the ring gear 944 is static (beingfixedly mounted via the flex mount 947 to the engine static structure919), while the planet carrier 946 and the plurality of planet gears 942therein, rotate about the centerline axis 912. The planet carrier 946constrains the plurality of planet gears 942 such that the plurality ofplanet gears 942 rotate together around the sun gear 940, while alsoenabling each planet gear of the plurality of planet gears 942 to rotateabout its own axis. The rotation of the planet carrier 946 is conveyedto the fan through the fan shaft 948.

FIG. 9B illustrates the mounting assembly 900 of FIG. 9A translated intoa representative vibratory system where each of the flex coupling 945,the flex mount 947, and the fan frame 949 are shown by representativestructural properties of the members, the representative structuralproperties being the structural stiffness (K) and the damping (C) of therespective members of the mounting assembly 900. As shown, each of theflex coupling 945, the flex mount 947, and the fan frame 949 includesthe representative structural properties (structural stiffness anddamping) in each of the lateral direction, the bending direction, andthe torsional direction.

For example, FIG. 9B represents the gearbox supporting structure interms of structural properties characterizing the nature of the couplingbetween the gearbox assembly 938 and the flex coupling 945. The flexcoupling 945 may be represented in terms of a flex coupling lateralstiffness K_(fc) ^(L), a flex coupling bending stiffness KA, a flexcoupling torsional stiffness K_(fc) ^(T), a flex coupling lateraldamping C_(fc) ^(L), a flex coupling bending damping C_(fc) ^(B), and aflex coupling torsional damping C_(fc) ^(T). In this way, the flexcoupling 945 acts as a dashpot damper that mechanically dampensvibrations through the flex coupling 945 from the gearbox assembly 938and/or from the low-speed shaft 936 (e.g., the midshaft).

FIG. 9B represents the gearbox supporting structure in terms ofstructural properties characterizing the nature of the coupling betweenthe gearbox assembly 938 and the flex mount 947. The flex mount 947 maybe represented in terms of a flex mount lateral stiffness K_(fm) ^(L), aflex mount bending stiffness K_(fm) ^(B), a flex mount torsionalstiffness K_(fm) ^(T), a flex mount lateral damping C_(fm) ^(L), a flexmount bending damping C_(fm) ^(B), and a flex mount torsional dampingC_(fm) ^(T). In this way, the flex mount 947 acts as a dashpot damperthat mechanically dampens vibrations through the flex mount 947 from thegearbox assembly 938 to the engine static structure 919.

FIG. 9B represents the gearbox supporting structure in terms ofstructural properties characterizing the nature of the coupling betweenthe gearbox assembly 938 and the fan frame 949. The fan frame 949 may berepresented in terms of a fan frame lateral stiffness K_(ff) ^(L), a fanframe bending stiffness K_(ff) ^(B), a fan frame torsional stiffnessK_(ff) ^(T), a fan frame lateral damping C_(ff) ^(L), a fan framebending damping C_(ff) ^(B), and a fan frame torsional damping C_(ff)^(T). In this way, the fan frame 949 acts as a dashpot damper thatmechanically dampens vibrations through the fan frame 949 from thegearbox assembly 938 to the engine static structure 919.

The gearbox mounting systems and configurations in FIGS. 8A and 9A canbe translated into a representative vibratory system, as shown in FIGS.8B and 9B, respectively. Each interface to the gear box, whether a fanframe, flex mount, or flex coupling has geometric qualities thattranslate to lateral, bending, and torsional stiffness and dampingelements. For example, the flex mount support system may have relativelythin-walled undulating supports engineered to possess specific valuesfor stiffness and damping. Support wall thickness and support memberspan or extent play a critical role in determining stiffness and dampingvalues. Thinner members certainly allow for lower values stiffnessquantities and shorter spans or member lengths contribute to highervalues stiffness properties. Similarly, the two flex mount flex elementson the input shaft use member thickness and outer diameter to controlstiffness and damping. As member thickness decreases and diaphragmdiameter increases, stiffness properties decrease in the mountinglocation. For the fan frame support, it is good practice to design thismounting element and location to be as stiff as possible whileminimizing weight. The fan support frame needs a high degree ofstiffness due to potential fan overloads that can occur; like in a bladeout failure scenario. Therefore, the design approach for the flex mountand flex element lateral and bending stiffness values are desired to benotably softer than the fan support frame, which allows for the gearboxsystem to follow the fan frame support movement while generating lowreaction forces and moments at the flex mount and flex coupling mountinglocations. Conversely, the torsional stiffness of the flex mount andflex coupling mounting elements is desired to be design as stiff aspossible since these elements are in the main torque transmission torquepath with the fan.

The flex coupling, the flex mount, and the fan frame of FIGS. 8A and 9Apermit the gearbox assembly (e.g., the gearbox assembly 838 and/or thegearbox assembly 938) to move to absorb bending moments applied by thefan shaft and/or the low-speed shaft. For example, the stiffness (K) andthe damping (C) (e.g., the bending stiffness, the lateral stiffness, andthe torsional stiffness, as well as the bending damping, the lateraldamping, and the torsional bending) of the gearbox assembly (e.g., thegearbox assembly 838 of FIGS. 8A and 8B and/or the gearbox assembly of938 of FIGS. 9A and 9B), can be selected to absorb the bending momentssuch that the effects of the vibratory response of the gearbox assemblyon the low-speed shaft can be tuned such that the vibrations of thegearbox assembly during operation do not excite the low-speed shaft atthe first-order bending mode when the low-speed shaft is operating at,or near, its critical speed, as detailed further below.

FIG. 10 illustrates an enlarged, schematic side view of a gearboxassembly 1038 with a mounting assembly 1000 for a gas turbine engine1010, taken at a centerline axis 1012 of the gas turbine engine 1010. Alow-speed shaft 1036 is shown schematically in FIG. 10 and extendsbeyond the cut view of FIG. 10 to illustrate the length L_(MSR) of thelow-speed shaft 1036. The mounting assembly 1000 is that for a planetaryconfiguration, as described with respect to FIGS. 9A and 9B. That is, aring gear 1044 is coupled with a flex mount 1047 to an engine staticstructure 1019. The plurality of planet gears 1042 is constrained withina planet carrier 1046, which is coupled to a fan shaft 1048, and a sungear 1040 is coupled with a flex coupling 1045 to the low-speed shaft1036. Although not shown in FIG. 10 , the fan shaft 1048 may be coupledwith a fan frame to the engine static structure 1019, such as describedwith respect to FIGS. 9A and 9B.

The gearbox assembly 1038 can include an oil transfer device 1050. Theoil transfer device 1050 allows an oil flow of lubricant (e.g., oil) toflow into the gearbox assembly 1038 and to lubricate the plurality ofplanet gears 1042, which in turn lubricates the sun gear 1040 and thering gear 1044. Although shown with respect to a planetaryconfiguration, the oil transfer device 1050 may be provided in a gearboxassembly having a star configuration (e.g., as shown and described withrespect to FIGS. 8A and 8B).

FIGS. 11A to 11C illustrate degrees of freedom associated withstructural stiffness K and damping coefficient C. These degrees offreedom characterize the most significant directions of movementaffecting the respective stiffness or damping properties of thecomponent as it interacts with the gearbox and engine frame(s)supporting it under loading conditions. The structural stiffness K andthe damping coefficient C representations allowed the inventors toquantify the structural dynamic behavior of these degrees of freedom ina sufficiently accurate and representative manner, accounting for allfactors in the component design that effects load transmission into thegearbox, thereby effecting the length L_(MSR) of the low-speed shaft.

In FIGS. 11A to 11C, the Z-axis coincides with the centerline axis ofthe gas turbine engine (as shown in FIGS. 1 to 3 ) and extends in theaxial direction A (FIGS. 1 to 3 ), the Y-axis extends perpendicular tothe Z-axis in the radial direction R (FIGS. 1 to 3 ), and the X-axisextends perpendicular to the Z-axis in the circumferential direction C(e.g., into and out of the page in FIGS. 1 to 3 ).

In FIG. 11A, the lateral stiffness K^(L) and the lateral damping C^(L)affect the lateral stiffness and the lateral damping of the respectivemounting component (e.g., the flex mount, the fan frame, and the flexcoupling). This results in the lateral stiffness K^(L) and the lateraldamping C^(L) affecting the movement of the respective component in thelateral direction. The lateral direction includes the linear motion ofthe component in a Y-axis radial direction 1100 and an X-axiscircumferential direction 1110.

In FIG. 11B, the bending stiffness K^(B) and the bending damping C^(B)affect the bending stiffness and the bending damping of the respectivemounting component (e.g., the flex mount, the fan frame, and the flexcoupling). This results in the bending stiffness K^(B) and the bendingdamping C^(B) affecting the rotational movement of the respectivecomponent in the bending direction. The bending direction includes thebending or rotational motion of the component in a yaw direction 1120and a pitch direction 1130.

In FIG. 11C, the torsional stiffness K^(T) and the torsional dampingC^(T) affect the torsional stiffness and the torsional damping of therespective mounting component (e.g., the flex mount, the fan frame, andthe flex coupling). This results in the torsional stiffness K^(T) andthe torsional damping C^(T) affecting the rotational movement of therespective component in a torsional direction 1140 about the enginecenterline (e.g., about the centerline axis or Z-axis). This representsthe load path of the gears and the torque of the respective componentwith respect to the fan (e.g., any of the fans or fan assembliesdetailed herein).

FIG. 12A illustrates an enlarged, schematic cross-sectional side view ofa gearbox assembly 1238 with a mounting assembly 1200 for a gas turbineengine 1210, taken at a centerline axis 1212 of the gas turbine engine1210. The gearbox assembly 1238 can be utilized as any of the gearboxassemblies detailed herein and the gas turbine engine 1210 can be any ofthe gas turbine engines detailed herein. The mounting assembly 1200shown is that for a star configuration gearbox, as detailed above. Thegearbox assembly 1238 includes a planet carrier 1250 that carries orhouses a plurality of planet gears 1252. A ring gear 1256 is connectedto a fan shaft 1239 as shown by arrow 1244. The plurality of planetgears 1252 rotate within the ring gear 1256 and are driven by a sun gear1254 that is connected to a low-speed shaft 1224 as shown by arrow 1246.The rotation of the low-speed shaft 1224 is transmitted through theplanet gears 1252 to the ring gear 1256 which rotates the fan shaft1239, and, thus, rotates a fan (e.g., any of the fans or fan assembliesdetailed herein).

The planet carrier 1250 includes a carrier arm 1258 that is connected toa first end of a flex mount 1260 at a connection 1286 (FIG. 12B). Thefirst end of the flex mount 1260 includes a deflection limiter 1261, asshown in more detail in FIG. 12B. The flex mount 1260 is connected at asecond end to a torque cone 1262 at a connection 1282 (FIG. 12A). Thetorque cone 1262 and the first end of the flex mount 1260 are connectedto each other and an engine static structure 1248 at a connection 1284(FIG. 12B). The torque cone 1262 includes a generally frusto-conicalsection between a first end at the connection 1282 and a second end atthe connection 1284. The design of the torque cone and the flex mountare only one embodiment of the supporting structure for the planetcarrier. The mounting system can include a different design and formbased on application requirements.

FIG. 12B shows an enlarged, schematic, cross-sectional side view of aportion of the mounting assembly 1200 at detail 1285 in FIG. 12A.Referring to FIG. 12B, the deflection limiter 1261 includes an outer rim1288 that is stationary due to its connection 1284 to the engine staticstructure 1248 and an inner rim 1290 that can oscillate or vibrate atthe connection 1286 to the carrier arm 1258. For example, the outer rim1288 and the inner rim 1290 extend axially and are spaced from eachother such that a gap 1268 is formed between the outer rim 1288 and theinner rim 1290. The gap 1268 has a clearance 1280 (FIG. 12D) thatdefines a size of the gap 1268 in the radial direction. The outer rim1288 and the inner rim 1290 are annular.

During operation, the gearbox assembly 1238 vibrates as the gears rotateand torque is transferred from the low-speed shaft 1224 to the fan shaft1239 through the gearbox assembly 1238. For example, the fan shaft 1239can apply bending moments through the gearbox assembly 1238, and theflex mount 1260 can absorb the bending moments. The vibration may causethe gearbox assembly 1238 to deflect in the axial direction (e.g.,lateral deflection), the radial direction (e.g., radial deflection),and/or the circumferential direction (e.g., torsional deflection). Thedeflection limiter 1261 prevents the gearbox assembly 1238 from beingdisplaced or being deflected beyond a threshold level. For example, asthe gearbox assembly 1238 deflects in the radial direction, the innerrim 1290 moves radially towards the outer rim 1288. As the gearboxassembly 1238 continues to move in the radial direction, the inner rim1290 will continue to move towards the outer rim 1288 until the innerrim 1290 contacts the outer rim 1288, thereby preventing the gearboxassembly 1238 from being deflected beyond the threshold level. Thethreshold level can be adjusted by adjusting the clearance 1280 (FIG.12D) of the gap 1268. For example, the threshold level is decreased bysetting the inner rim 1290 closer to the outer rim 1288 (e.g.,decreasing the size of the gap 1268), and the threshold level isincreased by setting the inner rim 1290 further away from the outer rim1288 (e.g., increasing the size of the gap 1268). Thus, the deflectionlimiter 1261 can be tuned to achieve a desired maximum deflection of thegearbox assembly 1238. The deflection limiter 1261 can also similarlylimit deflections in the axial and circumferential directions, asdetailed further below with respect to FIG. 12E. In this way, thedeflection limiter 1261 can be tuned to prevent the deflections of thegearbox assembly 1238 from exciting the low-speed shaft, therebyaffecting the MSR.

In some embodiments, the first end of the flex mount 1260 includes anoil feed passage 1266 for feeding oil (as shown by arrow 1270) into thegap 1268, also referred to as a damper land in such embodiments. In suchembodiments, the deflection limiter 1261 also functions as a damper,also referred to as a squeeze film damper (SFD). The oil is provided tothe gap 1268 through the oil feed passage 1266. The gap 1268 provides acontrol volume for the damper and includes the clearance 1280 (FIG.12D), also referred to as a squeeze film damper clearance in suchembodiments. The carrier arm 1258 includes an end seal 1276, and an endseal 1264 is connected to the flex mount 1260 and the carrier arm 1258by the connection 1286. The end seals 1264, 1276 provide end sealclearances 1274 through which oil may expelled (as shown by arrows 1272)as the damper dissipates vibratory energy. The end seals 1264, 1276 canalso limit deflections of the gearbox assembly 1238 in the axialdirection. In embodiments that do not include a SFD, the end seals 1264,1276 can be walls that prevent deflections of the gearbox assembly 1238in the axial direction.

Referring to FIGS. 12C and 12D, the integral damper formed by the outerrim 1288 and the inner rim 1290 of the flex mount 1260 includes springs1278 integrally formed in the flex mount 1260. Each spring 1278 includesa radial flex element 1204 and a torsional flex element 1206. In thisway, the springs 1278 help to damper vibrations in both the radialdirection and the circumferential direction. The springs 1278 can beintegrally formed in the flex mount 1260 by, for example,electric-discharge machining (EDM). The springs 1278 are providedcircumferentially around the flex mount 1260. A squeeze film damper(SFD) segment 1208 is defined between each circumferentially spaced setof springs 1278. A stop 1201 is provided between SFD segments 1208 toreduce or limit excursions of the mounting assembly 1200 with respect tothe engine static structure 1248.

Referring to FIGS. 12B and 12D, oil is fed into the gap 1268 which isbounded by the end seals 1264, 1276 on either side of the gap 1268. Theend seal clearance 1274 (FIG. 12B) controls the damping levels in thedamper in combination with a feed orifice diameter 1202 to the gap 1268.A significant vibration mode of the gearbox assembly 1238 (FIG. 12A) maybe excited by three different rotating elements: 1) the low-speed shaft1224 (FIG. 12A) imbalance; 2) the planet gear 1252 (FIG. 12A) imbalance;and 3) the fan shaft 1239 (FIG. 12A) imbalance. Other engine loads canalso be transmitted to the gearbox assembly 1238 such as oscillatingaerodynamic loads from the fan, transient maneuver loads, and torsionalexcitations for the gear train. The deflection limiter 1261 can mitigateor reduce these vibrations and will also reduce or minimize the dynamicforces transmitted to the engine structure.

FIG. 12E is an enlarged, schematic side view of the mounting assembly1200, taken at detail 1287 in FIG. 12C. FIG. 12E shows the deflectionlimiter 1261 includes one or more teeth 1295 that extend in thecircumferential direction, extend in the radial direction, and extendsin the axial direction (e.g., into the page in the view shown in FIG.12E). The deflection limiter 1261 also includes one or more gaps 1289through which the one or more teeth 1295 are disposed. The one or moregaps 1289 include a radially extending wall 1291 and a circumferentiallyextending wall 1293. While the one or more teeth 1295 are shown asextending from the inner rim 1290, the one or more teeth 1295 can extendfrom the outer rim 1288. Similarly, the one or more gaps 1289 are shownin the outer rim 1288 in FIG. 12E, but can be formed in the inner rim1290.

During operation, the gearbox assembly 1238 (FIG. 12A) deflects in theradial direction, the axial direction, and/or the circumferentialdirection, as detailed above. When the gearbox assembly 1238 deflectsradially, the inner rim 1290 and one or more teeth 1295 of thedeflection limiter 1261 deflect toward the circumferentially extendingwall 1293, and contact the circumferentially extending wall 1293 tolimit the deflection in the radially direction. Similarly, when thegearbox assembly 1238 deflects circumferentially, the inner rim 1290 andthe one or more teeth 1295 deflect toward the radially extending wall1291, and contact the radially extending wall 1291 to limit thedeflection in the circumferential direction. The one or more teeth 1295can similarly limit deflection in the axial direction when the gearboxassembly 1238 deflects axially.

FIG. 13 is an enlarged, schematic, cross-sectional side view of agearbox assembly 1338 with a mounting assembly 1300 for a gas turbineengine 1310, taken at a centerline axis 1312 of the gas turbine engine1310, according to another embodiment. The gearbox assembly 1338 can beutilized as any of the gearbox assemblies detailed herein and the gasturbine engine 1310 can be any of the gas turbine engines detailedherein. The mounting assembly 1300 shown is that for a planetaryconfiguration gearbox, as detailed. The gearbox assembly 1338 includes aplanet carrier 1350 that carries or houses a plurality of planet gears1352. In the planetary configuration, the planet carrier 1350 isconnected to a fan shaft 1339 as shown by arrow 1344. A ring gear 1356is stationary in the planetary configuration, and a flex mount 1360 ofthe mounting assembly 1300 is connected to the ring gear 1356. Theplurality of planet gears 1352 rotate within the ring gear 1356 and aredriven by a sun gear 1354 that is connected to a low-speed shaft 1324 asshown by arrow 1346. The rotation of the low-speed shaft 1324 istransmitted through the planet gears 1352 to the planet carrier 1350which rotates the fan shaft 1339, and, thus, rotates a fan (e.g., any ofthe fans or fan assemblies detailed herein). The mounting assembly 1300also includes a torque cone 1362 and a deflection limiter 1361 thatcoupled the gearbox assembly 1338 to an engine static structure 1348, asdetailed above.

FIG. 14 is an enlarged, schematic, partial cross-sectional side view ofa mounting assembly 1400 for a gearbox assembly of a gas turbine engine1410, taken at a centerline axis of the gas turbine engine, according tothe present disclosure. The mounting assembly 1400 includes a flex mount1460 and a torque cone 1462 and can be coupled to a planet carrier or toa ring gear of the gearbox assembly for coupling the gearbox assembly toan engine static structure 1448. The mounting assembly 1400 includes adeflection limiter 1461 including an outer rim 1488 that is stationarydue to a connection 1484 to the engine static structure 1448 and aninner rim 1490 that can oscillate or vibrate at a connection 1486 to acarrier arm 1458. For example, the outer rim 1488 and the inner rim 1490extend radially and are spaced from each other such that a gap 1468 isformed between the outer rim 1488 and the inner rim 1490. The gap 1468has a clearance 1480 that defines a size of the gap 1468 in the axialdirection. In this way, the deflection limiter 1461 operates similarlyto the deflection limiter 1261 of FIG. 12B, but the inner rim 1490deflects axially towards the outer rim 1488 to prevent deflections inthe axial direction.

The deflection limiter 1461 can also function as a damper similar to thedeflection limiter 1261 (FIG. 12B). For example, the deflection limiter1461 can include a first end seal 1463 and a second end seal 1465 andthe gap 1468 can receive oil (as shown by arrow 1470) therein and expeloil (as shown by arrow 1472) therefrom such that a squeeze film damperis provided.

FIG. 15 is an enlarged, schematic, partial cross-sectional view of aring gear assembly 1502, according to the present disclosure. The ringgear assembly 1502 is depicted as a dual ring gear/damper housing, alsoreferred to as a ring gear 1504. The ring gear 1504 can be utilized asany of the ring gears detailed herein. The ring gear assembly 1502includes a damper 1505 having a damper housing 1506. The ring gear 1504is connected to the damper housing 1506. The ring gear 1504 includes ageared wall 1508 that includes a plurality of ring gear teeth 1510. Theplurality of ring gear teeth 1510 are bihelical gear teeth. A detaileddescription of bihelical gear teeth is provided below. The plurality ofring gear teeth 1510 can include any type of gear teeth, such as, forexample, spur gear teeth, helical gear teeth, or the like. The ring gear1504 also includes a damper housing engagement wall 1512 extendingradially outward from the geared wall 1508. The ring gear 1504 alsoincludes a flexible damping wall 1514 connected to the geared wall 1508,and a damper housing attachment member 1516. The damper housingattachment member 1516 includes a longitudinal wall 1518 radially spacedapart from the geared wall 1508 and extending in an axial direction fromthe flexible damping wall 1514. The longitudinal wall 1518 can include aplurality of ring gear scavenge openings 1520 extending therethrough toallow a lubricant (e.g., oil) to flow therethrough. The damper housingattachment member 1516 further includes a damper housing connecting wall1522 connected with the longitudinal wall 1518 and extending outwardtherefrom in a radial direction. The damper housing connecting wall 1522includes a plurality of connecting openings 1524 therethrough forconnecting the ring gear 1504 to the damper housing 1506.

The flexible damping wall 1514 is shown as being a generallysemi-circular tubular-shaped wall that extends circumferentially, andthat connects to the geared wall 1508 and to the longitudinal wall 1518of the damper housing attachment member 1516. The flexible damping wall1514 has a flexible damping wall thickness 1526. The generallysemi-circular tubular-shape of the flexible damping wall 1514 providesfor radial flexure and damping of the geared wall 1508 during operationof the gearbox assembly, and the flexible damping wall thickness 1526 isa parameter provides a predetermined radial stiffness based on radialloads anticipated to be imparted to the geared wall 1508 by the planetgears (e.g., any of the planet gears detailed herein) during operationof the gearbox assembly.

The damper housing 1506 includes a damper housing radial wall 1528extending in the radial direction and extending circumferentially. Thedamper housing radial wall 1528 includes a plurality of frame mountingopenings 1530 for mounting the ring gear assembly 1502 to an enginestatic structure (e.g., any of the engine static structures detailedherein) via a plurality of fasteners (omitted for clarity). In this way,the ring gear assembly 1502 is utilized in a gearbox assembly having aplanetary configuration (e.g., a rotating planet carrier while the ringgear 1504 remains stationary). The damper housing radial wall 1528 alsoincludes a plurality of ring gear mounting openings 1532 for connectingthe ring gear 1504 to the damper housing 1506 via fasteners (omitted forclarity) provided through the connecting openings 1524 of the ring gear1504 and the ring gear mounting openings 1532.

The damper housing 1506 further includes a damper housing ring gearengagement wall 1534 extending radially inward from the damper housingradial wall 1528 and extending in the axial direction and in thecircumferential direction. The damper housing ring gear engagement wall1534 engages the longitudinal wall 1518 so as to provide support to thelongitudinal wall 1518 in a radially inward direction.

The ring gear assembly 1502 also includes a deflection limiter 1540 thatincludes a gap 1542. The deflection limiter 1540 functions substantiallysimilarly as the deflection limiter 1261 of FIGS. 12A to 12D to limitdeflections of the ring gear 1504 in the axial direction, the radialdirection, and/or in the circumferential direction. In some embodiments,the deflection limiter 1540 can also include and operate as an integraldamper (e.g., a squeeze film damper).

Components inside of the gearbox assembly, such as the gears or thebearings, can also affect the vibrational response of the gearboxassembly. FIG. 16A is a schematic, cross-sectional side view of agearbox assembly 1646 for a gas turbine engine, taken along a centerlineaxis 1612 of the gas turbine engine, according to the presentdisclosure. The gearbox assembly 1646 can be utilized as any of thegearbox assemblies and in any of the gas turbine engines detailedherein, but is particularly useful for gearbox assemblies having a gearratio of more than 4:1, or of 4:1 to 12:1, or of 7:1 to 12:1, or of 4:1and 10:1, or of 5:1 and 8:1. The gearbox assembly 1646 in FIG. 16A issuited for use with gas turbine engines that have large fan blades, suchas the open fan configuration of FIG. 3 . The gearbox assembly 1646 inFIG. 16A is one embodiment of the gearbox 355 in FIG. 3 .

The gearbox assembly 1646 includes an epicyclic gear assembly 1647 in acompound symmetrical arrangement. The epicyclic gear assembly 1647includes a sun gear 1652, a plurality of planet gears 1654 (only one ofwhich is visible in FIG. 16A), and a ring gear 1656. For clarity, only aportion of the gears is shown. The gearbox assembly 1646 is a star type(e.g., star configuration) or a rotating ring gear type gearbox assembly(e.g., the ring gear 1656 is rotating and a planet carrier 1680 is fixedand stationary). In such an arrangement, the fan (e.g., any of the fansor fan assemblies detailed herein) is driven by the ring gear 1656. Inthis way, the ring gear 1656 is an output of the gearbox assembly 1646.However, other suitable types of gearbox assemblies may be employed. Inone non-limiting embodiment, the gearbox assembly 1646 may be aplanetary arrangement, in which the ring gear 1656 is held fixed, withthe planet carrier 1680 allowed to rotate. In such an arrangement, thefan is driven by the planet carrier 1680. In this way, the plurality ofplanet gears 1654 are the output of the gearbox assembly 1646.

An input shaft 1636 is coupled to the sun gear 1652. The input shaft1636 is coupled to the power turbine section of the gas turbine engine(e.g., via a low-speed shaft). The input shaft 1636 is coupled to thelow-speed shaft, having midshaft length L_(MSR), as detailed above. Thering gear 1656 is coupled via an output shaft 1644 to the fan androtates to drive rotation of the fan about the centerline axis 1612. Forexample, the output shaft 1644 is coupled to a fan shaft (e.g., any ofthe fan shafts detailed herein) of the fan. In some embodiments, theoutput shaft 1644 and the fan shaft are formed as a single integralcomponent.

Each of the plurality of planet gears 1654 is a compound gear thatincludes a first stage planet gear 1660 and a second stage planet gear1662 coupled together. The first stage planet gear 1660 includes agreater diameter than a diameter of the second stage planet gear 1662.In some embodiments, the diameter of the first stage planet gear 1660 isequal to or less than the diameter of the second stage planet gear 1662.The diameters of the first stage planet gear 1660 and the second stageplanet gear 1662 can be selected to change a gear ratio split betweenthe first stage planet gear 1660 and the second stage planet gear 1662.Each of the sun gear 1652, the plurality of planet gears 1654, and thering gear 1656 comprises teeth about their periphery to intermesh withthe other gears. For example, each of the sun gear 1652, the pluralityof planet gears 1654, and the ring gear 1656 are bihelical gears withfirst and second sets of helical teeth that are all inclined at the sameacute angle relative to a planet gear axis. The helical teeth of aplanet gear 1654 are further detailed below with respect to FIG. 16B.The sun gear 1652 comprises a first set of sun gear teeth 1664 and asecond set of sun gear teeth 1666. Each of the first stage planet gears1660 includes a first set of planet gear teeth 1668 and a second set ofplanet gear teeth 1670, and each of the second stage planet gears 1662includes a third set of planet gear teeth 1672 and a fourth set ofplanet gear teeth 1674. The ring gear 1656 includes a first set of ringgear teeth 1676 and a second set of ring gear teeth 1678. The sun gear1652, the plurality of planet gears 1654, and the ring gear 1656 mayinclude any type of gear, such as, for example, spur gears (e.g., gearteeth that are straight cut and are not set at an angle relative to theplanet gear axis), or the like.

The first set of planet gear teeth 1668 and the second set of planetgear teeth 1670 of the first stage planet gear 1660 mesh with the firstset of sun gear teeth 1664 and the second set of sun gear teeth 1666 ofthe sun gear 1652, respectively. The third set of planet gear teeth 1672of the second stage planet gear 1662 meshes with the first set of ringgear teeth 1676 of the ring gear 1656. The fourth set of planet gearteeth 1674 of the second stage planet gear 1662 meshes with the secondset of ring gear teeth 1678 of the ring gear 1656.

Each of the planet gears 1654 of the plurality of planet gears 1654includes a pin 1682 about which a respective planet gear 1654 rotates.The pin 1682 is coupled to the planet carrier 1680 and is disposedwithin a bore 1683 of a respective planet gear 1654. Lubricant (e.g.,oil) is provided between the pin 1682 and a respective planet gear 1654such that the planet gear 1654 rotates with respect to the pin 1682. Thesecond stage planet gear 1662 is supported by one or more rollerbearings 1684 that are disposed within the bore 1683. FIG. 16A shows theone or more roller bearings 1684 include two roller bearings 1684including a first roller bearing 1684 a and a second roller bearing 1684b. A respective planet gear 1654, however, can include any number ofroller bearings 1684, as desired. The second roller bearing 1684 b islocated aft of the first roller bearing 1684 a. In this way, the firstroller bearing 1684 a is referred to as a forward roller bearing and thesecond roller bearing 1684 b is referred to as an aft roller bearing.The roller bearings 1684 allow rotation of the planet gear 1654 withrespect to the pin 1682.

The gearbox assembly 1646 includes a gear ratio that defines a ratio ofthe speed of the input gear (e.g., the sun gear 1652) to the speed ofthe output (e.g., the ring gear 1656) through the gearbox assembly 1646.Embodiments of the present disclosure detailed herein provide forincreased gear ratios for a fixed gear envelope (e.g., with the samesize ring gear), or alternatively, a smaller diameter ring gear may beused to achieve the same gear ratios. Thus, the embodiments disclosedherein allow for gear ratios suitable for large diameter engines, or forsmaller diameter engines. A total gear ratio of the planet gear 1654includes a first gear ratio of the first stage planet gear 1660 and asecond gear ratio of the second stage planet gear 1662. The first gearratio of the first stage planet gear 1660 is less than the second gearratio of the second stage planet gear 1662. In some embodiments, thefirst gear ratio of the first stage planet gear 1660 is greater than orequal to the second gear ratio of the second stage planet gear 1662. Inthe embodiment of FIG. 16A, the total gear ratio of the planet gear 1654is between seven (7:1) and twelve (12:1). The total gear ratio isselected based on engine size and power requirements and a selection ofcomponents for a particular gearbox assembly 1646. For example, thetotal gear ratio is based on the speed of the fan (e.g., the tip speedof the fan) and the speed of the low-pressure turbine (e.g., based onthe number of stages of the low-pressure turbine).

FIG. 16B is a schematic view of a planet gear 1654 with a first stageplanet gear 1660 and a second stage planet gear 1662, according to thepresent disclosure. While the planet gear is described herein, thefollowing description is applicable to any of the gears of the gearboxassembly 1646 (FIG. 16A). The planet gear 1654 is a bihelical gear, alsoreferred to as a double helical gear, in which the gear teeth arearranged in a “herringbone” pattern. Because each of the gear meshes(sun-to-planet and planet-to-ring) has a bihelical gear tooth profile,there is no relative movement possible parallel to an axis 1613 betweenthe planet gear 1654 and the sun gear or the ring gear, or in otherwords there is no axial compliance between these elements. In this way,such an arrangement of the gear teeth provides for a stiffer meshbetween the gears of the gearbox assembly as compared to gear teethhaving a spur type (e.g., parallel with the axis 1613) arrangement.

Each planet gear tooth of the first set of planet gear teeth 1668 andthe second set of planet gear teeth 1670 defines a first helix axis 1667along a length of each respective gear tooth of the first stage planetgear 1660. The first set of planet gear teeth 1668 are mirrored withrespect to the second set of planet gear teeth 1670. The first helixaxis 1667 is normal to an end face of each respective gear tooth of thefirst stage planet gear 1660. The first helix axis 1667 is disposed at afirst helix angle β₁ with respect to the axis 1613 (e.g., with respectto the axis of rotation of the gear). The greater the first helix angleβ₁, the greater the stiffness of the mesh between the gears. However,increasing the first helix angle B1 is limited due to geometry and heattreatment of the gear teeth, tooling, and/or the resulting axial load.Thus, the first helix angle β₁ is in a range of twenty-two point fivedegrees (22.5°) to thirty-two point five degrees (32.5°) to balance therequirement for an increased stiffness in the mesh with the aboveconsiderations of the geometry and heat treatment, the tooling, and/orthe resulting axial load. Preferably, the first helix angle β₁ is thirtydegrees (30°) to provide a greater stiffness in the mesh, whileaccounting for the considerations above for gearboxes with gear ratiosabove 7:1. The first helix angle β₁ can be changed to change a stiffnessof the mesh between the gears.

Each planet gear tooth of the third set of planet gear teeth 1672 andthe second set of planet gear teeth 1674 of the second stage planet gear1662 defines a second helix axis 1669 along a length of each respectivegear tooth of the second stage planet gear 1662. The third set of planetgear teeth 1672 are mirrored with respect to the fourth set of planetgear teeth 1674. The second helix axis 1669 is normal to an end face ofeach respective gear tooth of the second stage planet gear 1662. Thesecond helix axis 1669 is disposed at a second helix angle β₂ withrespect to the axis 1613 (e.g., with respect to the axis of rotation ofthe gear). The second helix angle β₂ is in a range of twenty-two pointfive degrees (22.5°) to thirty-two point five degrees (32.5°).Preferably, the first helix angle β₂ is thirty degrees (30°), asdetailed above. The second helix angle β₂ can be changed to change astiffness of the mesh between the gears. The second helix angle β₂ isthe same as the first helix angle β₁ to balance loads on the planet gear1654 during operation. In some embodiments, the second helix angle β₂ isdifferent than the first helix angle β₁.

The stiffness of the gears in the gearbox assembly affect the vibrationsof the gearbox assembly, thereby effecting vibrations from the gearboxassembly to the low-speed shaft, and, thus, the length L_(MSR) of thelow-speed shaft and the MSR. For example, vibrations in the gearboxassembly (e.g., from the gears) propagate through the gearbox assembly,thereby vibrating the gearbox assembly, and the vibrations can propagatethrough to the low-speed shaft. Accordingly, the gear teeth type and anangle (e.g., the helix angle) of the gear teeth can be selected toachieve a particular vibrational response in the gearbox assembly so asto avoid exciting the low-speed shaft when the low-speed shaft isoperating subcritical, as detailed further below.

FIG. 17 is a schematic view of a planet gear 1754 having a single stagefor a gearbox assembly (e.g., any of the gearbox assemblies detailedherein), according to the present disclosure. The single stage gear canbe used in gearbox assemblies for gas turbine engines having lower gearratios (e.g., less than 7:1). While the planet gear is described herein,the following description is applicable to any of the gears of thegearbox assembly. The planet gear 1754 is a bihelical gear, similar tothe planet gear 1654 of FIG. 16B. The planet gear 1754 includes a firstset of planet gear teeth 1768 and a second set of planet gear teeth1770. The first set of planet gear teeth 1768 are mirrored with respectto the second set of planet gear teeth 1770 such that the gear teeth arein a “herringbone” pattern.

Each planet gear tooth of the first set of planet gear teeth 1768 andthe second set of planet gear teeth 1770 defines a helix axis 1767 alonga length of each respective gear tooth of the planet gear 1754. Thehelix axis 1767 is normal to an end face of each respective gear toothof the planet gear 1754. The helix axis 1767 is disposed at a helixangle R with respect to an axis 1713 (e.g., with respect to the axis ofrotation of the gear) of the planet gear 1754. The helix angle R is in arange of twenty-two point five degrees (22.5°) to thirty-two point fivedegrees (32.5°). Preferably, the helix angle R is twenty-five degrees(25°) to increase the stiffness of the mesh while accounting for theconsiderations above for gearboxes with a gear ratio less than 7:1. Thehelix angle R can be changed to change a stiffness of the mesh betweenthe gears. The helix angle R can be changed to change a stiffness of thegears, and, therefore, a vibratory response in the gearbox assembly, asdetailed above.

FIG. 18 illustrates an enlarged, schematic side cross-sectional view ofa gearbox assembly 1838 with a mounting assembly 1800 for a gas turbineengine 1810, taken at a centerline axis 1812 of the gas turbine engine1810. The gearbox assembly 1838 can be utilized as any of the gearboxassemblies and the gas turbine engine 1810 can be any of the gas turbineengines detailed herein. The gearbox assembly 1838 is substantiallysimilar to the gearbox assembly 938 of FIGS. 9A and 9B and includes aplanetary configuration. For example, the gearbox assembly 1838 includesa sun gear 1840, a plurality of planet gears 1842, a ring gear 1844, alow-speed shaft 1836 coupled to the sun gear 1840. The sun gear 1840 iscoupled via a flex coupling 1845 to the low-speed shaft 1836. Theplurality of planet gears 1842 are coupled together by a planet carrier1846. In the embodiment of FIG. 18 , the planet carrier 1846 is coupled,via a fan shaft 1848, to a fan (e.g., any of the fans or fan assembliesdetailed herein) to drive rotation of the fan about the centerline axis1812. The fan shaft 1848 is coupled to a fan frame 1849 via a fanbearing 1850. The ring gear 1844 is coupled via a flex mount 1847 to anengine static structure 1819. The flex coupling 1845, the flex mount1847, and the fan frame 1849 define the mounting assembly 1800 for thegearbox assembly 1838. As described herein, the flex coupling 1845, theflex mount 1847, and the fan frame 1849 may be referred to as mountingmembers.

In FIG. 18 , the flex coupling 1845 is part of an input shaft 1851 thatextends from a forward bearing 1852 of the low-speed shaft 1836 to thesun gear 1840 (e.g., to an axially center of the sun gear 1840). Theflex coupling 1845 is also referred to as a decoupler, and includes oneor more flex plates 1854 that absorb and reduce deflections andvibrations from propagating from the gearbox assembly 1838 to thelow-speed shaft 1836 or from the low-speed shaft 1836 to the gearboxassembly 1838. In the embodiment shown in FIG. 18 , the one or more flexplates 1854 include a first flex plate 1854 a and a second flex plate1854 b spaced axially from each other along the input. The one or moreflex plates 1854 can include any number of flex plates located at anyaxial position along the input, as desired. The flex plates 1854 areintegral with the flex coupling 1845 and include axial gaps that absorbthe deflections in an axial direction so that propagation of thedeflections through the flex coupling 1845 is reduced. Accordingly, theflex coupling 1845 can be tuned or can be changed to achieve aparticular desired vibrational frequency response such that vibrationsof the gearbox assembly 1838 do not excite the low-speed shaft 1836 whenthe redline speed is subcritical.

The input shaft 1851 includes an input shaft length L_(input) thatextends axially from the forward bearing 1852 to the sun gear 1840(e.g., an axial center of the sun gear 1840). The input shaft lengthL_(input) is equal to an aft decoupler length L_(dplr_aft), a decouplerlength L_(dcplr), and a forward decoupler length L_(dcplr_fwd), addedtogether. The aft decoupler length L_(dplr_aft) extends from the forwardbearing 1852 to the first flex plate 1854 a, the decoupler lengthL_(dcplr) extends from the first flex plate 1854 a to the second flexplate 1854 b, and the forward flex length L_(dcplr_fwd) extends from thesecond flex plate 1854 b to the sun gear 1840 (e.g., to an axiallycenter of the sun gear 1840). The flex coupling 1845 also includes adecoupler height H_(dcplr) and one or more decoupler radii. Thedecoupler height is a height of the flex plates 1854 in the radialdirection from the input shaft 1851. The one or more decoupler radii isan inner radius of the input shaft 1851. The one or more decoupler radiiinclude a first decoupler radius R_(dcplr1) and a second decouplerradius R_(dcplr2). In the embodiment of FIG. 18 , the first decouplerradius R_(dcplr1) is equal to the second decoupler radius R_(dcplr2)such that the input shaft 1851 has a constant inner radius. In someembodiments the first decoupler radius R_(dcplr1) is different than thesecond decoupler radius R_(dcplr2) such that the input shaft 1851 has avariable inner radius (e.g., the inner radius of the input shaft 1851changes along the axial direction).

In consideration of midshaft operating speeds, whether during anaircraft maximum thrust at takeoff, redline or cruise operatingcondition, it is desirable to have any anticipated dynamic loading ofthe gearbox caused by midshaft motion to not act as to amplify or excitefundamental or principle mode(s) of the gearbox through the sungear—midshaft coupling. It is also desirable to avoid a dynamicexcitation communicated through the sun gear/midshaft coupling andinfluenced by modal characteristics (represented generally by modelfrequency FGBx) of the gearbox assembly to act as to excite fundamentalmode(s) of the midshaft. To achieve this end result, it is desirable tohave a decoupler moment stiffness KM_(dcplr) of the flex coupling 1845and a decoupler shear stiffness KS_(dcplr) of the flex coupling 1845(e.g., a moment stiffness and a shear stiffness at the sun gear-midshaftcoupling) being such as to neither cause significant excitation of afundamental midshaft mode, nor a dynamic excitation from the midshaftcommunicated at this coupling to cause significant excitation of afundamental mode of the gearbox assembly. The decoupler moment stiffnessKM_(dcplr) is an overturning moment stiffness of the flex coupling 1845(e.g., a torque of the flex coupling 1845 applied radially on the flexcoupling 1845), including the decoupler moment stiffness of the firstflex plate 1854 a and the decoupler moment stiffness of the second flexplate 1854 b. The decoupler shear stiffness KS_(dcplr) is a stiffness ofthe flex coupling 1845 (e.g., between the first flex plate 1854 a andthe second flex plate 1854 b) in the axial direction. The stiffness ofthe flex coupling 1845 (e.g., the decoupler moment stiffness KM_(dcplr)and the decoupler shear stiffness KS_(dcplr)) should be selected so asto not amplify midshaft properties or so as not excite the gearboxassembly 1838 by midshaft dynamic behavior during engine operation.

Various rig tests and measurements taken to simulate engine operationalconditions, accounting for any differences between a dynamic responsefor a recently fielded engine and an engine after several operationalcycles, revealed common patterns in dynamic behavior formidshaft-gearbox interactions to inform the design of the flex coupling1845 to avoid the modal coupling between gearbox and midshaft explainedabove. It was found that a decoupler moment stiffness KM_(dcplr) of theflex coupling 1845 in a range of 50 klb*in/rad to 200 klb*in/rad, and adecoupler shear stiffness KS_(dcplr) of the flex coupling 1845 in arange of 100 klb/in to 500 klb/in, should substantially avoidintolerable or sustained dynamic amplification of the gearbox assembly1838 or the midshaft (e.g., the low-speed shaft) when there isexcitation of either the gearbox assembly 1838 or the midshaft duringengine operations. In this way, the flex coupling 1845 prevents thegearbox assembly 1838 from dynamically exciting the midshaft, andprevents the midshaft from dynamically exciting the gearbox assembly1838. The decoupler moment stiffness KM_(dcplr) of the flex coupling1845 is expressed in klb*in/rad, and the decoupler shear stiffnessKS_(dcplr) of the flex coupling 1845 is expressed in klb/in. In view ofthe foregoing, the decoupler moment stiffness KM_(dcplr) of the flexcoupling 1845 and the decoupler shear stiffness KS_(dcplr) of the flexcoupling 1845 are desired to satisfy the relationships (2) and (3),respectively:

$\begin{matrix}{{KM}_{dcplr} = \frac{E*K_{m}*R_{dcplr}^{4}}{H_{dcplr}}} & (2)\end{matrix}$ $\begin{matrix}{{KS}_{dcplr} = \frac{E*K_{m}*R_{dcplr}^{4}}{L_{dcplr}^{2}}} & (3)\end{matrix}$

In the relationships (2) and (3), E is the Young's modulus for theparticular material of the flex coupling 1845 expressed in lb/in², K_(m)is a correction factor for the decoupler moment stiffness KM_(dcplr) andfor the decoupler shear moment KS_(dcplr) to account for variousdifferent materials that the flex coupling 1845 could be made of,R_(dcplr) is the decoupler radius (e.g., the first decoupler radiusR_(dcplr1) and/or the second decoupler radius R_(dcplr2)) expressed ininches, H_(dcplr) is the decoupler height (e.g., of the one or more flexplates 1854) expressed in inches, and L_(dcplr) is a length of the flexcoupling 1845 (e.g., an axial length that extends from the first flexplate 1854 a to the second flex plate 1854 b) expressed in inches. Thematerial of flex coupling 1845 can include, for example, metal alloys,titanium, steel, or the like. The K_(m) correction factor is a constantand is in a range of 0.13×10⁻³ to 0.27×10⁻³.

The decoupler moment stiffness is a function of an aft decoupler momentstiffness KM_(dcplr_aft) of the first flex plate 1854 a and/or a forwarddecoupler moment stiffness KM_(dcplr_fwd) of the second flex plate 1854b. The aft decoupler moment stiffness KM_(dcplr_aft) is approximatelyequal to the forward decoupler moment stiffness KM_(dcplr_fwd). In someembodiments, the aft decoupler moment stiffness KMaft is different thanthe forward decoupler moment stiffness KMfwd. For example, the forwarddecoupler moment stiffness KM_(dcplr_fwd) can be two to three (2 to 3)times the aft decoupler moment stiffness KM_(dcplr_aft). In someembodiments, the aft decoupler moment stiffness KM_(dcplr_aft) is two tothree (2 to 3) times the forward decoupler moment stiffnessKM_(dcplr_fwd).

The stiffness of the flex coupling 1845 can be modified to producesignificant changes in the critical speed. Embodiments are listed inTABLE 6. These embodiments were considered as possible designs thatcould modify the effective stiffness of the flex coupling 1845 in such away to be compatible with engine architecture and without requiringmodifications or limitations on the targeted operating range for asubcritical shaft.

TABLE 6 Redline ODR L_(MSR) D_(MSR) E/rho Teff Mode L_(MSR)/ Speed MSREmb. Description (in) (in) (in⁻¹) (in) (RPM) D_(MSR) (ft/sec)(ft/sec)^(1/2) 33 4-bearing 60.6 2.80 1.30E+08 0.32 9208 22 111 231system, including bottle boring, with a gearbox having 200 klb/ineffective shear stiffness 34 4-bearing 60.6 2.80 1.30E+08 0.32 10238 22123 244 system, including bottle boring, with a gearbox having 500klb/in effective shear stiffness 35 4-bearing 60.6 2.80 1.30E+08 0.329467 22 114 235 system, including bottle boring, with a gearbox having200 klb/in effective shear stiffness and 50 klb*in/rad effective momentstiffness 36 4-bearing 60.6 2.80 1.30E+08 0.32 11587 22 140 259 system,including bottle boring, with a gearbox having 500 klb/in effectiveshear stiffness and 200 klb*in/rad effective moment stiffness

Embodiments 33, 34, 35, and 36 all use a steel alloy materialcomposition. Embodiments 33 to 36 all use a four-bearing system, and allhave a convex thickness profile similar to the example of FIG. 6C,having been manufactured with a bottle boring manufacturing technique.These embodiments include the same bearing arrangement, but differ fromeach other in terms of stiffness (e.g., decoupler shear stiffness and/ordecoupler moment stiffness). The values shown in TABLE 6 illustrate thatembodiments 34 and 36 achieve a greater stiffness (e.g., the decouplershear stiffness and/or the decoupler moment stiffness) than embodiments33 and 35, respectively, generally resulting in substantially higherredline speeds and correspondingly higher MSR.

To further achieve the end result of avoiding a dynamic excitationcommunicated through the sun gear-midshaft coupling, it is desirable tohave an effective gearbox assembly stiffness K_(GBX) (representing thecombined stiffness properties of the gearbox couplings and mounts, thedeflection limiters, dampers, and the gears of the gearbox assembly inthe lateral direction, the bending direction, and/or the torsionaldirection, as detailed above) and mass M_(GBX) representation of thegearbox assembly at the sun-gear midshaft coupling being such as toneither cause significant excitation of a fundamental midshaft mode, nora dynamic excitation from the midshaft communicated at this coupling tocause significant excitation of a fundamental mode of the gearboxassembly, wherein the associated modal frequency of the gearbox assemblymode at the sun gear—midshaft coupling is represented by the term(K_(GBX)/M_(GBX))^(1/2). The fundamental modal properties of the gearboxassembly, FGBX, should not therefore act as to amplify midshaftproperties or be excitable by midshaft dynamic behavior during engineoperation.

Various rig tests and measurements taken to simulate engine operationalconditions, accounting for any differences between a dynamic responsefor a recently fielded engine and an engine after several operationalcycles, revealed common patterns in dynamic behavior formidshaft-gearbox interactions to inform the design of gearbox assembliesto avoid the modal coupling between gearbox and midshaft explainedabove. It was found that a gearbox assembly mode FGBX less than ninetyfive percent (95%), or greater than one hundred five percent (105%) ofthe midshaft mode F_(MIDSHAFT) (i.e., the bending mode associated withthe midshaft critical speed) should substantially avoid intolerable orsustained dynamic amplification of gearbox or midshaft primary modes(i.e., FGBX, F_(MIDSHAFT) respectively) when there is excitation ofeither these modes during engine operations. In this way, the gearboxassembly mode FGBX is prevented from affecting the midshaft modeF_(MIDSHAFT). The redline speed for a midshaft as expressed in Hertz is(MSR/(L_(MSR)/D_(MSR))){circumflex over ( )}2*12/(πD_(MSR)), whereL_(MSR) is the length L of the midshaft and D_(MSR) is the outerdiameter D of the midshaft. The gearbox assembly mode FGBx in Hertz, inradians/sec as it influences the mid-shaft behavior, or the midshaftaffecting gearbox dynamic behavior, is

${\sqrt{\frac{K_{GBX}}{M_{GBX}}}/2\pi},$where K_(GBX) is the gearbox assembly stiffness and M_(GBX) is thegearbox assembly mass representation at the sun gear—midshaft coupling.In view of the foregoing, the desired gearbox assembly modal propertiesat the sun gear—midshaft coupling are desired to satisfy either of therelationships (4) and (5):

$\begin{matrix}{{11.4{\left( {{MSR}/\left( {L_{MSR}/D_{MSR}} \right)} \right)\hat{}2}/\left( {\pi D_{MSR}} \right)} > {\sqrt{\frac{K_{GBX}}{M_{GBX}}}/2\pi}} & (4)\end{matrix}$ $\begin{matrix}{{12.6{\left( {{MSR}/\left( {L_{MSR}/D_{MSR}} \right)} \right)\hat{}2}/\left( {\pi D_{MSR}} \right)} < {\sqrt{\frac{K_{GBX}}{M_{GBX}}}/2\pi}} & (5)\end{matrix}$

FIG. 19A illustrates examples of ranges and/or values for a midshaftrating, with respect to OD speed at redline. The plot indicates valuesfor the midshaft rating (MSR). Specifically, FIG. 19A shows a range 1915defined by MSR between 200 (ft/sec)⁻¹ (curve 1920) and 300 (ft/sec)⁻¹(curve 1925), for redline speeds from fifty to two hundred and fiftyfeet per second.

FIG. 19B illustrates ranges and/or values for a midshaft rating, withrespect to L/D ratio. The plot indicates values for the midshaft rating(MSR). Specifically, FIG. 19B shows a range 1930 defined by MSR between200 (ft/sec)⁻¹ (curve 1935) and 300 (ft/sec)⁻¹ (curve 1940), for L/Dratios from twelve to thirty-seven.

Further aspects of the present disclosure are provided by the subjectmatter of the following clauses.

A turbomachine engine includes a core engine having one or morecompressor sections, one or more turbine sections that includes a powerturbine, and a combustion chamber in flow communication with thecompressor sections and turbine sections. The turbomachine engine alsoincludes a shaft that is coupled to the power turbine and that ischaracterized by a midshaft rating (MSR) between two hundred(ft/sec)^(1/2) and three hundred (ft/sec)^(1/2), and a redline speedbetween fifty and two hundred fifty feet per second (ft/sec).

A turbomachine engine includes a core engine having one or morecompressor sections, one or more turbine sections that includes a powerturbine, and a combustion chamber in flow communication with thecompressor sections and turbine sections. The turbomachine engine alsoincludes a shaft that is coupled to the power turbine and that ischaracterized by a midshaft rating (MSR) between two hundred(ft/sec)^(1/2) and three hundred (ft/sec)^(1/2), a length L, an outerdiameter D, and a ratio of L/D between twelve and thirty-seven.

The turbomachine engine of any preceding clause, wherein theturbomachine engine is configured to operate up to a redline speedwithout passing through a critical speed associated with a first-orderbending mode of the shaft.

The turbomachine engine of any preceding clause, wherein theturbomachine engine is configured to operate up to the redline speedwithout passing through a critical speed associated with a first-orderbending mode of the shaft.

The turbomachine engine of any preceding clause, wherein the MSR isbetween one hundred ninety (ft/sec)^(1/2) and two hundred forty-five(ft/sec)^(1/2).

The turbomachine engine of any preceding clause, wherein the MSR isbetween two hundred twelve (ft/sec)^(1/2) and two hundred sixty(ft/sec)^(1/2).

The turbomachine engine of any preceding clause, wherein the MSR isbetween two hundred forty-seven (ft/sec)^(1/2) and two hundred ninety(ft/sec)^(1/2).

The turbomachine engine of any preceding clause, wherein the redlinespeed is between sixty and ninety ft/sec.

The turbomachine engine of any preceding clause, wherein the redlinespeed is between seventy-five and one hundred seventy-five ft/sec.

The turbomachine engine of any preceding clause, wherein the redlinespeed is between one hundred fifty and two hundred fifty ft/sec.

The turbomachine engine of any preceding clause, wherein the ratio ofL/D is between twelve and twenty.

The turbomachine engine of any preceding clause, wherein the ratio ofLID is between sixteen and thirty.

The turbomachine engine of any preceding clause, wherein the ratio ofLID is between twenty-six and thirty-seven.

The turbomachine engine of any preceding clause, wherein the shaft is acomposite shaft made of at least two different materials.

The turbomachine engine of any preceding clause, wherein the shaft has alength L and a reduced mass density at a midpoint along the length L.

The turbomachine engine of any preceding clause, wherein the shaft has areduced mass density at a midpoint along the length L.

The turbomachine engine of any preceding clause, wherein the shaft is ahollow convex shaft with a reduced wall thickness at the midpoint, avariable inner diameter, and a constant outer diameter.

The turbomachine engine of any preceding clause, wherein the shaft is ahollow convex shaft with a reduced wall thickness at the midpoint, aconstant inner diameter, and a variable outer diameter.

The turbomachine engine of any preceding clause, wherein the shaft iscoupled to the power turbine at a first mounting point, and wherein theshaft is also coupled to one of the compressor sections at a secondmounting point.

The turbomachine engine of any preceding clause, wherein the shaft issupported by at least a first bearing and a second bearing.

The turbomachine engine of any preceding clause, wherein the shaft has alength L that is measured as the distance between the first bearing andthe second bearing.

The turbomachine engine of any preceding clause, wherein the length L ismeasured as the distance between the first bearing and the secondbearing.

The turbomachine engine of any preceding clause, wherein at least onebearing is a duplex bearing that has an overturning moment capability.

The turbomachine engine of any preceding clause, wherein each bearing isone of a ball bearing and a roller bearing.

The turbomachine engine of any preceding clause, wherein the firstbearing is positioned between the first mounting point and the secondmounting point, and wherein the second mounting point is positionedbetween the first bearing and the second bearing.

The turbomachine engine of any preceding clause, wherein the firstbearing and the second bearing support the shaft in an inboundconfiguration in which the first bearing and the second bearing arepositioned between the first mounting point and the second mountingpoint.

The turbomachine engine of any preceding clause, wherein the firstbearing and the second bearing support the shaft in an outboundconfiguration in which the first mounting point and the second mountingpoint are positioned between the first bearing and the second bearing.

The turbomachine engine of any preceding clause, wherein the shaft isfurther supported by a third bearing and a fourth bearing.

The turbomachine engine of any preceding clause, wherein the firstbearing and the second bearing are a first pair of duplex bearings, andthe third bearing and the fourth bearing are a second pair of duplexbearings, wherein the first pair of duplex bearings and the second pairof duplex bearings support the shaft in an inbound overturning momentconfiguration in which the first pair of duplex bearings and the secondpair of duplex bearings are positioned between the first mounting pointand the second mounting point.

The turbomachine engine of any preceding clause, wherein the firstbearing and the second bearing are a first pair of duplex bearings, andthe third bearing and the fourth bearing are a second pair of duplexbearings, wherein the first pair of duplex bearings and the second pairof duplex bearings support the shaft in an outbound overturning momentconfiguration in which the first mounting point and the second mountingpoint are positioned between the first pair of duplex bearings and thesecond pair of duplex bearings.

The turbomachine engine of any preceding clause, wherein the firstbearing, the second bearing, the third bearing, and the fourth bearingsupport the shaft in a four-bearing straddle configuration in which thefirst bearing and the second bearing are positioned between the firstmounting point and the second mounting point, and the first mountingpoint and the second mounting point are positioned between the thirdbearing and the fourth bearing.

The turbomachine engine of any preceding clause, wherein the firstbearing, the second bearing, the third bearing, and the fourth bearingsupport the shaft in a four-bearing outbound configuration in which thefirst mounting point and the second mounting point are positionedbetween a first group of bearings comprising the first bearing and thesecond bearing, and a second group of bearings comprising the thirdbearing and the fourth bearing.

In another aspect, a method includes using a turbomachine engine with acore having one or more compressor sections, one or more turbinesections that includes a power turbine, and a combustion chamber in flowcommunication with the compressor sections and turbine sections.

The method also includes driving a shaft that is coupled to the powerturbine and that is characterized by a midshaft rating (MSR) between twohundred (ft/sec)^(1/2) and three hundred (ft/sec)^(1/2), and a redlinespeed between fifty and two hundred fifty feet per second (ft/sec).

In another aspect, a method includes using a turbomachine engine with acore having one or more compressor sections, one or more turbinesections that includes a power turbine, and a combustion chamber in flowcommunication with the compressor sections and turbine sections. Themethod also includes driving a shaft that is coupled to the powerturbine and that is characterized by a midshaft rating (MSR) between twohundred (ft/sec)^(1/2) and three hundred (ft/sec)^(1/2), a length L, anouter diameter D, and a ratio of L/D between twelve and thirty-seven.

The method of any preceding clause, wherein the turbomachine engine isconfigured to operate up to a redline speed without passing through acritical speed associated with a first-order bending mode of the shaft.

The method of any preceding clause, wherein the turbomachine engine isconfigured to operate up to the redline speed without passing through acritical speed associated with a first-order bending mode of the shaft.

The method of any preceding clause, wherein the MSR is between onehundred ninety (ft/sec)^(1/2) and two hundred forty-five (ft/sec)^(1/2).

The method of any preceding clause, wherein the MSR is between twohundred twelve (ft/sec)^(1/2) and two hundred sixty (ft/sec)^(1/2).

The method of any preceding clause, wherein the MSR is between twohundred forty-seven (ft/sec)^(1/2) and two hundred ninety(ft/sec)^(1/2).

The method of any preceding clause, wherein the redline speed is betweensixty and ninety ft/sec.

The method of any preceding clause, wherein the redline speed is betweenseventy-five and one hundred seventy-five ft/sec.

The method of any preceding clause, wherein the redline speed is betweenone hundred fifty and two hundred fifty ft/sec.

The method of any preceding clause, wherein the ratio of L/D is betweentwelve and twenty.

The method of any preceding clause, wherein the ratio of LID is betweensixteen and thirty.

The method of any preceding clause, wherein the ratio of L/D is betweentwenty-six and thirty-seven.

The method of any preceding clause, wherein the shaft is a compositeshaft made of at least two different materials.

The method of any preceding clause, wherein the shaft has a length L anda reduced mass density at a midpoint along the length L.

The method of any preceding clause, wherein the shaft has a reduced massdensity at a midpoint along the length L.

The method of any preceding clause, wherein the shaft is a hollow convexshaft with a reduced wall thickness at the midpoint, a variable innerdiameter, and a constant outer diameter.

The method of any preceding clause, wherein the shaft is a hollow convexshaft with a reduced wall thickness at the midpoint, a constant innerdiameter, and a variable outer diameter.

The method of any preceding clause, wherein the shaft is coupled to thepower turbine at a first mounting point, and wherein the shaft is alsocoupled to one of the compressor sections at a second mounting point.

The method of any preceding clause, wherein the shaft is supported by atleast a first bearing and a second bearing.

The method of any preceding clause, wherein the shaft has a length Lthat is measured as the distance between the first bearing and thesecond bearing.

The method of any preceding clause, wherein the length L is measured asthe distance between the first bearing and the second bearing.

The method of any preceding clause, wherein at least one bearing is aduplex bearing that has an overturning moment capability.

The method of any preceding clause, wherein each bearing is one of aball bearing and a roller bearing.

The method of any preceding clause, wherein the first bearing ispositioned between the first mounting point and the second mountingpoint, and wherein the second mounting point is positioned between thefirst bearing and the second bearing.

The method of any preceding clause, wherein the first bearing and thesecond bearing support the shaft in an inbound configuration in whichthe first bearing and the second bearing are positioned between thefirst mounting point and the second mounting point.

The method of any preceding clause, wherein the first bearing and thesecond bearing support the shaft in an outbound configuration in whichthe first mounting point and the second mounting point are positionedbetween the first bearing and the second bearing.

The method of any preceding clause, wherein the shaft is furthersupported by a third bearing and a fourth bearing.

The method of any preceding clause, wherein the first bearing and thesecond bearing are a first pair of duplex bearings, and the thirdbearing and the fourth bearing are a second pair of duplex bearings,wherein the first pair of duplex bearings and the second pair of duplexbearings support the shaft in an inbound overturning momentconfiguration in which the first pair of duplex bearings and the secondpair of duplex bearings are positioned between the first mounting pointand the second mounting point.

The method of any preceding clause, wherein the first bearing and thesecond bearing are a first pair of duplex bearings, and the thirdbearing and the fourth bearing are a second pair of duplex bearings,wherein the first pair of duplex bearings and the second pair of duplexbearings support the shaft in an outbound overturning momentconfiguration in which the first mounting point and the second mountingpoint are positioned between the first pair of duplex bearings and thesecond pair of duplex bearings.

The method of any preceding clause, wherein the first bearing, thesecond bearing, the third bearing, and the fourth bearing support theshaft in a four-bearing straddle configuration in which the firstbearing and the second bearing are positioned between the first mountingpoint and the second mounting point, and the first mounting point andthe second mounting point are positioned between the third bearing andthe fourth bearing.

The method of any preceding clause, wherein the first bearing, thesecond bearing, the third bearing, and the fourth bearing support theshaft in a four-bearing outbound configuration in which the firstmounting point and the second mounting point are positioned between afirst group of bearings comprising the first bearing and the secondbearing, and a second group of bearings comprising the third bearing andthe fourth bearing.

A turbomachine engine comprising a fan section having a fan shaft, acore engine including one or more compressor sections, one or moreturbine sections that includes a power turbine, and a combustion chamberin flow communication with the compressor sections and turbine sections,a low-speed shaft coupled to the power turbine and having a midshaftthat extends from a forward bearing to an aft bearing, the low-speedshaft being characterized by a midshaft rating (MSR) between two hundred(ft/sec)^(1/2) and three hundred (ft/sec)^(1/2), the low-speed shafthaving a redline speed between fifty and two hundred fifty feet persecond (ft/sec), and the turbomachine engine being configured to operateup to the redline speed without passing through a critical speedassociated with a first-order bending mode of the low-speed shaft, and agearbox assembly that couples the fan shaft to the low-speed shaft, thegearbox assembly being characterized by a gearbox assembly mode that isless than 95% of a midshaft mode of the midshaft or greater than 105% ofthe midshaft mode.

The turbomachine engine of the preceding clause, the MSR being betweenone hundred ninety (ft/sec)^(1/2) and two hundred forty-five(ft/sec)^(1/2).

The turbomachine engine of any preceding clause, the MSR being betweentwo hundred twelve (ft/sec)^(1/2) and two hundred sixty (ft/sec)^(1/2).

The turbomachine engine of any preceding clause, the MSR being betweentwo hundred forty-seven (ft/sec)^(1/2) and two hundred ninety(ft/sec)^(1/2).

The turbomachine engine of any preceding clause, further comprising aflex mount that couples the gearbox assembly to an engine staticstructure of the turbomachine engine.

The turbomachine engine of any preceding clause, further comprising afan frame that couples the fan shaft to an engine static structure, thefan frame connecting to the engine static structure forward of the oneor more compressor sections.

The turbomachine engine of any preceding clause, further comprising aflex coupling that couples the gearbox assembly to the low-speed shaft.

The turbomachine engine of any preceding clause, the flex coupling beingcharacterized by a decoupler moment stiffness in a range of 50klb*in/rad to 200 klb*in/rad.

The turbomachine engine of any preceding clause, the flex coupling beingcharacterized by a decoupler shear stiffness in a range of 100 klb/in to500 klb/in.

The turbomachine engine of any preceding clause, the gearbox assemblyincluding a deflection limiter that limits deflections of the gearboxassembly beyond a threshold level.

The turbomachine engine of any preceding clause, the deflection limiterlimiting deflections of the gearbox assembly in an axial direction, aradial direction, and/or a circumferential direction.

The turbomachine engine of any preceding clause, the deflection limiterincluding one or more teeth and one or more gaps having one or morewalls, the one or more teeth contacting the one or more walls to limitdeflections of the gearbox assembly.

The turbomachine engine of any preceding clause, the one or more teethextending in the circumferential direction, in the radial direction, andin the axial direction.

The turbomachine engine of any preceding clause, the one or more wallsincluding a radially extending wall and a circumferentially extendingwall.

The turbomachine engine of any preceding clause, the gearbox assemblyincluding one or more dampers that damper vibrations of the gearboxassembly.

The turbomachine engine of any preceding clause, the one or more dampersincluding a squeeze film damper.

The turbomachine engine of any preceding clause, the one or more dampersincluding a dashpot damper.

The turbomachine engine of any preceding clause, wherein the gearboxassembly including a plurality of gears including a sun gear, aplurality of planet gears, and a ring gear.

The turbomachine engine of any preceding clause, the plurality of gearsbeing helical gears.

The turbomachine engine of any preceding clause, the helical gearsincluding a helix angle in a range of 22.5° to 32.5°.

The turbomachine engine of any preceding clause, the decoupler momentstiffness being equal to

$\frac{E*K_{m}*R_{dcplr}^{4}}{H_{dcplr}},$E being a Young's modulus of a material of the flex coupling, K_(m)being a correction factor, R_(dcplr) being a decoupler radius of theflex coupling, and H_(dcplr) being a decoupler height of the flexcoupling.

The turbomachine engine of any preceding clause, K_(m) being in a rangeof 0.13×10⁻³ to 0.27×10⁻³.

The turbomachine engine of any preceding clause, the decoupler shearstiffness being equal to

$\frac{E*K_{m}*R_{dcplr}^{4}}{L_{dcplr}^{2}},$wherein E being a Young's modulus of a material of the flex coupling,K_(m) being a correction factor, R_(dcplr) being a decoupler radius ofthe flex coupling, and L_(dcplr) being a decoupler length of the flexcoupling.

The turbomachine engine of any preceding clause, K_(m) being in a rangeof 0.13×10⁻³ to 0.27×10⁻³.

The turbomachine engine of any preceding clause, the turbomachine enginebeing an unducted fan engine.

The turbomachine engine of any preceding clause, the turbomachine enginebeing a ducted fan engine.

The turbomachine engine of any preceding clause, the gearbox assemblyhaving a gear ratio of 4:1 to 12:1, or 7:1 to 12:1, or 4:1 to 10:1, or5:1 to 9:1, or 6:1 to 9:1.

The turbomachine engine of any preceding clause, the gearbox assemblyhaving a gear ratio of 3:1 to 4:1, 3:5 to 4:1, 3.25:1 to 3.5:1, or 4:1to 5:1.

The turbomachine engine of any preceding clause, the gearbox assemblyfurther including an oil transfer device that allows an oil flow oflubricant to flow into the gearbox assembly to lubricate the pluralityof gears.

The turbomachine engine of any preceding clause, the deflection limiterincluding an outer rim and an inner rim, the inner rim moving towardsthe outer rim as the gearbox assembly deflects.

The turbomachine engine of any preceding clause, the deflection limiterlocated at the flex mount.

The turbomachine engine of any preceding clause, the flex mount coupledto the planet carrier.

The turbomachine engine of any preceding clause, the flex mount coupledto the ring gear.

The turbomachine engine of any preceding clause, the flex mount beingcoupled to the engine static structure by a torque cone.

The turbomachine engine of any preceding clause, the flex mountincluding an end seal.

The turbomachine engine of any preceding clause, the damper furthercomprising one or more springs.

The turbomachine engine of any preceding clause, the one or more springsincluding a radial flex element that dampers vibrations in the radialdirection and a torsional flex element that dampens vibrations in thecircumferential direction.

The turbomachine engine of any preceding clause, the inner rim and theouter rim being arranged such that the inner rim deflects radiallytowards the outer rim.

The turbomachine engine of any preceding clause, the inner rim and theouter rim being arranged such that the inner rim deflects axiallytowards the outer rim.

The turbomachine engine of any preceding clause, the ring gear includinga damper having a flexible damping wall.

The turbomachine engine of any preceding clause, the plurality of planetgears including compound planet gears including a first stage planetgear and a second stage planet gear.

The turbomachine engine of any preceding clause, the first stage planetgear having a first helix angle in a range of 22.5° to 32.5°, and thesecond stage planet gear having a second helix angle in a range of 22.5°to 32.5°.

The turbomachine engine of any preceding clause, the plurality of planetgears including single stage planet gears.

The turbomachine engine of any preceding clause, the helix angle beingthirty degrees (30°).

The turbomachine engine of any preceding clause, the gearbox assemblybeing decoupled from the low-speed shaft.

The turbomachine engine of any preceding clause, the flex couplingincluding a one or more flex plates that absorb deflections in the axialdirection.

The turbomachine engine of any preceding clause, the one or more flexplates including a first flex plate and a second flex plate.

The turbomachine engine of any preceding clause, further comprising anelectric machine.

The turbomachine engine of any preceding clause, the electric machinebeing coupled to the low-pressure shaft.

The turbomachine engine of any preceding clause, the electric machinebeing located aft of the power turbine.

The turbomachine engine of any preceding clause, the electric machinebeing an electric motor.

The turbomachine engine of any preceding clause, the electric machinebeing an electric generator.

The turbomachine engine of any preceding clause, the electric machineincluding a rotor and a stator, the rotor rotating with respect to thestator.

The turbomachine engine of any preceding clause, the rotor being coupledto the low-pressure shaft.

The turbomachine engine of any preceding clause, the rotor rotating withthe low-pressure shaft.

The turbomachine engine of any preceding clause, the electric machinebeing a motor-generator.

A method includes using a turbomachine engine with a fan section havinga fan shaft, and a core engine having one or more compressor sections,one or more turbine sections that includes a power turbine, and acombustion chamber in flow communication with the compressor sectionsand turbine sections. The method also includes driving a low-speed shaftthat is coupled to the power turbine and that has a midshaft thatextends from a forward bearing to an aft bearing, and that ischaracterized by a midshaft rating (MSR) between two hundred(ft/sec)^(1/2) and three hundred (ft/sec)^(1/2), and a redline speedbetween fifty and two hundred fifty feet per second (ft/sec). The methodfurther comprises operating the low-speed shaft at a linear speed up tothe redline speed without passing through a critical speed associatedwith a first-order bending mode of the low-speed shaft, and thelow-speed shaft has a midshaft that extends from a forward bearing to anaft bearing. The turbomachine further includes a gearbox assembly thatcouples the fan shaft to the low-speed shaft, the gearbox assemblycharacterized by a gearbox assembly mode less than 95% of a midshaftmode of the midshaft or greater than 105% of the midshaft mode.

The method of the preceding clause, the turbomachine engine being theturbomachine engine of any preceding clause.

The method of any preceding clause, further comprising driving thelow-speed shaft with an electric machine.

The method of any preceding clause, further comprising driving thelow-speed shaft to generate electric power in an electric machine.

The method of any preceding clause, further comprising driving a rotorof the electric machine with the low-speed shaft.

Although the foregoing description is directed to certain embodiments,other variations and modifications will be apparent to those skilled inthe art, and may be made without departing from the spirit or scope ofthe disclosure. Moreover, features described in connection with oneembodiment may be used in conjunction with other embodiments, even ifnot explicitly stated above.

The invention claimed is:
 1. A turbomachine engine comprising: a fansection having a fan shaft; a core engine including one or morecompressor sections, one or more turbine sections that includes a powerturbine, and a combustion chamber in flow communication with thecompressor sections and turbine sections; a low-speed shaft coupled tothe power turbine and having a midshaft that extends from a forwardbearing to an aft bearing, the low-speed shaft being characterized by amidshaft rating (MSR) between two hundred (ft/sec)^(1/2) and threehundred (ft/sec)^(1/2), wherein the low-speed shaft has a redline speedbetween fifty and two hundred fifty feet per second (ft/sec), and theturbomachine engine is configured to operate up to the redline speedwithout passing through a critical speed associated with a first-orderbending mode of the low-speed shaft; and a gearbox assembly that couplesthe fan shaft to the low-speed shaft, the gearbox assembly characterizedby a gearbox assembly mode that is less than 95% of a midshaft mode ofthe midshaft or greater than 105% of the midshaft mode.
 2. Theturbomachine engine of claim 1, wherein the MSR is between two hundred(ft/sec)^(1/2) and two hundred forty-five (ft/sec)^(1/2).
 3. Theturbomachine engine of claim 1, wherein the MSR is between two hundredtwelve (ft/sec)^(1/2) and two hundred sixty (ft/sec)^(1/2).
 4. Theturbomachine engine of claim 1, wherein the MSR is between two hundredforty-seven (ft/sec)^(1/2) and two hundred ninety (ft/sec)^(1/2).
 5. Theturbomachine engine of claim 1, further comprising a flex mount thatcouples the gearbox assembly to an engine static structure of theturbomachine engine.
 6. The turbomachine engine of claim 1, furthercomprising a fan frame that couples the fan shaft to an engine staticstructure, wherein the fan frame connects to the engine static structureforward of the one or more compressor sections.
 7. The turbomachineengine of claim 1, further comprising a flex coupling that couples thegearbox assembly to the low-speed shaft.
 8. The turbomachine engine ofclaim 7, wherein the flex coupling is characterized by a decouplermoment stiffness in a range of 50 klb*in/rad to 200 klb*in/rad.
 9. Theturbomachine engine of claim 7, wherein the flex coupling ischaracterized by a decoupler shear stiffness in a range of 100 klb/in to500 klb/in.
 10. The turbomachine engine of claim 1, wherein the gearboxassembly includes a deflection limiter that limits deflections of thegearbox assembly beyond a threshold level.
 11. The turbomachine engineof claim 10, wherein the deflection limiter limits deflections of thegearbox assembly in an axial direction, a radial direction, and/or acircumferential direction.
 12. The turbomachine engine of claim 11,wherein the deflection limiter includes one or more teeth and one ormore gaps having one or more walls, the one or more teeth contacting theone or more walls to limit deflections of the gearbox assembly.
 13. Theturbomachine engine of claim 12, wherein the one or more teeth extend inthe circumferential direction, in the radial direction, and in the axialdirection.
 14. The turbomachine engine of claim 12, wherein the one ormore walls include a radially extending wall and a circumferentiallyextending wall.
 15. The turbomachine engine of claim 1, wherein thegearbox assembly includes one or more dampers that dampen vibrations ofthe gearbox assembly.
 16. The turbomachine engine of claim 15, whereinthe one or more dampers include a squeeze film damper.
 17. Theturbomachine engine of claim 15, wherein the one or more dampers includea dashpot damper.
 18. The turbomachine engine of claim 1, wherein thegearbox assembly includes a plurality of gears including a sun gear, aplurality of planet gears, and a ring gear.
 19. The turbomachine engineof claim 18, wherein the plurality of gears are helical gears.
 20. Theturbomachine engine of claim 19, wherein the helical gears include ahelix angle in a range of 22.5° to 32.5°.